DNEPR User's guide

DNEPR

This User’s Guide contains technical data, the use of which is mandatory for: 

 

 

evaluation of spacecraft/Dnepr-1 launch vehicle compatibility; and 

preparation of all technical and operational documentation regarding a spacecraft 

launch on Dnepr-1 launch vehicle. 

All questions on the issues associated with the operation of the Dnepr Space Launch 

System that were not addressed in this User’s Guide should be sent to the address below: 

P.O. Box 7, Moscow, 123022, Russian Federation 

Tel.: (+7 095) 745 7258 

Fax: (+7 095) 232 3485 

E-mail: info@kosmotras.ru 

Current information relating to the Dnepr Space Launch System, activities of International 

Space Company Kosmotras, performed and planned launches of Dnepr launch vehicle can 

be found on ISC Kosmotras web-site: 

http://www.kosmotras.ru 

Issue 2, November 2001 


2

Document Change Record 

Issue Number Description Date Approved 

Issue 2 

Dnepr SLS User’s Guide, completely 

revised 

November 2001 

Stanislav I. Us, 

Designer General, 

Dnepr Program 


3

Table of Contents 

1. Introduction 10 

2. International Space Company Kosmotras 12 

2.1 ISC Kosmotras Permits and Authorities 12 

2.2 Dnepr Program Management. Dnepr Team and Responsibilities 12 

3. Purpose, Composition and Principal Characteristics of Dnepr Space Launch 15 

System 

4. Dnepr-1 Launch Vehicle 17 

4.1 General Description 17 

4.2 Spacecraft Injection Accuracy 22 

4.3 Launch Vehicle Axes Definition 22 

4.4 Space Head Module 24 

4.4.1 Space Head Module Design 24 

4.4.2 Payload Envelope 26 

4.5 Launch Vehicle Flight Reliability 30 

5. Baikonur Cosmodrome 31 

6. Dnepr SLS Ground Infrastructure 35 

6.1 Elements and General Diagram of Ground Infrastructure 35 

6.2 SC and SHM Processing Facility 35 

6.3 Spacecraft Fuelling Station 39 

6.4 Launch Complex 40 

7. Baikonur Operations Flow 42 

7.1 Preparation of LV for Integration with SHM 43 

7.2 Independent SC Processing 43 

7.3 Preparation of SHM for Mating with SC 44 

7.4 SC / SHM Integration 44 

7.5 Transportation of SHM Containing SC by Transporter Erector to the 45 

Silo Launcher 

7.6 SHM/LV Integration and LV Preparation for Launch 46 

7.7 Pre-launch Operations and LV Launch 47 

8. SC/LV Interfaces 49 


4

8.1 Mechanical Interface 49 

8.2 Electrical Interface 51 

9. Spacecraft Environments 54 

9.1 Stiffness Criteria (Frequency Requirements) 54 

9.2 Quasi-static and Dynamic Loads 54 

9.3 Vibration Loads 54 

9.4 Shock Loads 57 

9.5 Acoustic Loads 57 

9.6 Temperature and Humidity Conditions and Thermal Effect on 

58 

Spacecraft 

9.7 Pressure Underneath LV Fairing 59 

9.8 Gas-dynamic Effect on Spacecraft 59 

9.9 SC/LV Electromagnetic Compatibility 60 

9.10 Spacecraft Tests Required to Meet Dnepr LV Launch Services 

62 

Requirements 

10. Ground Qualification Tests 62 

11. Telemetry and Tracking 64 

12. Analysis of Flight Results 65 

13. Range Safety 66 

14. Transportation of Spacecraft and Associated Equipment. Customs 

69 

Clearance 

14.1 Transportation of Dummy Spacecraft to Ukraine for Ground Testing 69 

14.2 Transportation of Spacecraft and Associated Equipment to Baikonur 69 

Cosmodrome 

15. Principles of Pricing Policy and Contractual Relations 75 

16. Design and Technical Documents to Be Submitted by Spacecraft Authority 73 

17. Documents to Be Submitted to ISC Kosmotras by Spacecraft Authority 75 

18. Launch Campaign Schedule 76 


5

List of Figures 

Figure. 1-1 Dnepr-1 Lifts off from Baikonur Cosmodrome. September 26,2000 11 

Figure 2.2-1 Dnepr Program Operations Diagram 14 

Figure 4.1-1 Dnepr-1 General View 18 

Figure 4.1-2 SS-18 1 st and 2 nd Stages Inside TLC 18 

Figure 4.1-3 Dnepr-1 Performance Curves for Circular Orbits 19 

Figure 4.1-4 Mission Profile of Dnepr-1 LV Carrying a Large Spacecraft 20 

Figure 4.1-5 Mission Profile of Dnepr-1 LV with a Group of Spacecraft 21 

Figure 4.3-1 Launch Vehicle Major Axes 23 

Figure 4.4.1-1 SHM Configuration 1 – Standard Length 25 

Figure 4.4.1-2 SHM Configuration 2 – Standard Length 25 

Figure 4.4.1-3 SHM Configuration 1 – Extended by 850 mm 25 

Figure 4.4.1-4 SHM with 2-tier Layout 25 

Figure 4.4.2-1 Payload Envelope Available within SHM Configuration 1 with 27 

Standard Adapter and GDS 

Figure 4.4.2-2 Payload Envelope Available within SHM Configuration 2 with 28 

Standard Adapter and EPM 

Figure 4.4.2-3 Payload Envelope Available within SHM Configuration 1 Extended 29 

by 850 mm with Standard Adapter and GDS 

Figure 5-1 Yubileiniy Airfield 31 

Figure 5-2 Syrdarya River 31 

Figure 5-3 Cosmodrome Residential Area – Town of Baikonur 31 

Figure 5-4 Kosmonavt Hotel 32 

Figure 5-5 Baikonur Hotel 32 

Figure 5-6 Facilities of Sputnik Hotel 32 

Figure 6.1-1 Dnepr SLS at Baikonur Cosmodrome 35 

Figure 6.2-1 Layout of AITB, Site 42 36 

Figure 6.2-2 Layout of AITB, Site 31 37 

Figure 6.4-1 Dnepr Launch Complex Facilities (Sites 106 and 109) 40 

Figure 6.4-2 Transporter Erector 41 

Figure 7-1 Spacecraft Processing and Launch Schedule 42 


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Figure 7.1-1 Loading of TLC Containing LV 1 st and 2 nd Stages into Silo 

43 

Launcher 

Figure 7.2-1 Off-loading of Spacecraft Container from the Plane 43 

Figure 7.2-2 Spacecraft Processing 44 

Figure 7.4-1 SHM Integration 45 

Figure 7.5-1 SHM Loading into Transporter Erector 45 

Figure 7.6-1 SHM/LV Integration Inside Silo Launcher 46 

Figure 7.6-2 SHM Loading into Silo Launcher 47 

Figure 7.7-1 Dnepr LV Liftoff 48 

Figure 8.1-1 Typical Design of Spacecraft/Standard Adapter Attachment Point 50 

Figure 8.2-1 Schematics of SC/SC COE Transit Electrical Links 52 

Figure 9.7-1 Pressure Change Rate inside Payload Envelope 59 

Figure 11-1 Tracking Station 64 

Figure 12-1 Flight Data Processing Center 65 

Figure 18-1 Launch Campaign Schedule 76 


7

List of Tables 

Table 4.1-1 Dnepr-1 Main Characteristics 18 

Table 4.2-1 Spacecraft Injection Accuracy 22 

Table 5-1 Typical Routes of Transportation at Baikonur Cosmodrome 33 

Table 6.2-1 Characteristics of Area A, B and C, AITB at Site 31 38 

Table 9.2-1 Accelerations at SC/LV Interface during Transportation 55 

Table 9.2-2 Maximum Quasi-static and Dynamic Accelerations at SC/LV 

55 

Interface 

Table 9.3-1 Amplitude of Harmonic Oscillations at SC/LV Interface. 

56 

Longitudinal Axis (X) 

Table 9.3-2 Amplitude of Harmonic Oscillations at SC/LV Interface. Lateral 

56 

Axes (Y, Z) 

Table 9.3-3 Spectral Density of Vibro-accelerations at SC/LV Interface 56 

Table 9.4-1 Shock Spectrum at Spacecraft Attachment Points 57 

Table 9.5-1 Acoustic Loads 58 

Table 9.8-1 Maximum Values of 3 rd Stage Motor Plume Parameters Affecting 60 

Spacecraft 

Table 9.9-1 Maximum Residual Levels of RF Emissions 61 

Table 9.9-2 Maximum Levels of Electromagnetic Emissions 61 


8

Abbreviations 

AITB - Assembly, Integration and Test Building 

COE - Checkout Equipment 

ECOE - Electrical Checkout Equipment 

EPM - Encapsulated Payload Module 

FSUE - Federal State Unitary Enterprise 

GDS - Gas-dynamic Shield 

GPE - Ground Processing Equipment 

GPS - Global Positioning System 

ICBM - Intercontinental Ballistic Missile 

ICD - Interface Control Document 

ISC - International Space Company 

LCC - Launch Control Center 

LSA - Launch Services Agreement 

LSS - Launch Services Specifications 

LV - Launch Vehilce 

MoD - Ministry of Defense 

MoU - Memorandum of Understanding 

NSAU - National Space Agency of Ukraine 

RASA - Russian Aviation and Space Agency 

SC - Spacecraft 

SDB - State Design Bureau 

SHM - Space Head Module 

SLS - Space Launch System 

TLC - Transport and Launch Canister 

UHV - Ultra High Frequency 


9

1. Introduction 

Dnepr Space Launch System (SLS) is 

designed for expedient high accuracy 

injection of various single or multiple 

spacecraft (SC) weighing up to 3.7 metric 

tons into 300 – 900 km low earth orbits 

inclined 50.5, 64.5, 87.3 or 98 degrees. 

The core of the Dnepr SLS is the world’s 

most powerful intercontinental ballistic 

missile (ICBM) SS-18 or Satan, which 

possesses high performance 

characteristics and mission reliability 

confirmed through 159 launches (including 

2 launches under the Dnepr Program). 

The SS-18’s design and invariance of its 

control system allowed for creating, on its 

basis, a high-performance Dnepr launch 

vehicle (LV) equipped with a space head 

module (SHM), which meets modern 

requirements for SC injection means. 

The actual availability of a considerable 

number of missiles (about 150) that have a 

long operational life, of ground 

infrastructure at Baikonur Cosmodrome 

that comprises processing facilities and 

launch complex, drop zones and ground 

data processing complex, as well as of a 

team of the companies that developed the 

Dnepr SLS, ensures stable provision of 

launch services. 

Dnepr Program implementation is 

exercised in full compliance with the 

provisions of the treaties on reduction and 

limitation of the strategic offensive arms 

(START Treaties). 

As a first practical step in the Dnepr 

Program implementation, ISC Kosmotras 

jointly with the Russian Ministry of 

Defense prepared and launched the 

British Surrey Satellite Technology Limited 

UoSAT-12 satellite on a modified SS-18 

(Dnepr-1) rocket on April 21, 1999. 

The following step of the Dnepr Program 

evolution is mastering cluster launches of 

satellites belonging to different customers. 

September 26, 2000 marked the Dnepr 

launch with five spacecraft: MegSat-1 

(MegSat s.P.a., Italy), UniSat (La 

Sapienza University, Rome, Italy), 

SaudisSat-1A and 1B (Space Research 

Institute, Saudi Arabia) and TiungSat-1 

(ATSB, Malaysia). 

The User’s Guide contains information on 

basic characteristics, performance, initial 

data for spacecraft integration with the 

launch vehicle as well as spacecraft 

environments within the Dnepr SLS that 

will facilitate the preliminary evaluation of 

feasibility to utilize the Dnepr SLS for the 

launch of a specific spacecraft. 


10

Figure. 1-1 Dnepr-1 Lifts off from Baikonur Cosmodrome. September 26,2000 


11

2. International Space Company 

Kosmotras 

2.1 ISC Kosmotras Permits and 

Authorities 

International Space Company (ISC) 

Kosmotras (a joint stock company) was 

established in 1997 by aerospace 

agencies of Russia and Ukraine. 

ISC Kosmotras performs the development 

and commercial operation of the Dnepr 

SLS in accordance with the Russian 

Government Decree # 1156 “On Creation 

of Dnepr Space Launch System” dated 

October 5, 1998 and Ukrainian Cabinet of 

Ministers decree # 1246 “On Measures 

Regarding the Establishment of ISC 

Kosmotras” dated November 6, 1997. 

The activities on Dnepr SLS were 

recognized in the joint statement of the 

Russian and Ukrainian presidents issued 

on May 31, 1997, and on February 27, 

1998 were incorporated into the Program 

of Cooperation between Russia and 

Ukraine in the field of research and 

exploitation of outer space for peaceful 

purposes, and into the Russian Federal 

Space Program and National Space 

Program of Ukraine. 

In the joint statement of the Russian and 

Ukrainian presidents on cooperation in the 

field of rocket and space technology dated 

12th February 2001, the expansion of 

commercial application of Dnepr launch 

vehicle was quoted as top priority for long 

term cooperation of the two countries in 

the field of space related activities. 

In the Program of Cooperation between 

the Russian Aviation and Space Agency 

and National Space Agency of Ukraine in 

the field of research and exploitation of 

outer space for peaceful purposes for 

2001, which is an integral part of the long 

term cooperation program between Russia 

and Ukraine for 1998 - 2007, the Program 

of development and operation of the 

Dnepr Space Launch System was called 

as one of the priority projects for 2001. 

2.2 Dnepr Program Management. Dnepr 

Team and Responsibilities 

The top management body of ISC 

Kosmotras is its Board of Directors 

composed of the members from Russia, 

Ukraine and the Republic of Kazakhstan - 

heads of companies incorporated in ISC 

Kosmotras, officials of governmental 

entities, as well as leading specialists of 

the program. 

Direct Dnepr Program management is 

exercised by ISC Kosmotras principal 

office located in Moscow, Russia. 

Given below are Dnepr team members: 

Russia 

 

Russian Aviation and Space Agency 

(Moscow) - state support and 

supervision, provision of facilities and 

services at Baikonur Cosmodrome; 

Russian Ministry of Defense – 

allocation of SS-18 assets to be 

converted into Dnepr-1 launch 

vehicles, SS-18 storage and Dnepr-1 

standard launch operations; 

Joint Stock Company ASKOND 

(Moscow) – primary entity for Dnepr 

Program management; 


12

Joint Stock Company Rosobschemash 

Corporation (Moscow) - coordination of 

SS-18 elimination programs; 

FSUE Design Bureau of Special 

Machine Building (St. Petersburg) - 

primary entity for maintenance of 

launch complex and processing 

facilities; 

 

FSUE Central Scientific and Research 

Institute of Machine Building (Moscow) 

- scientific and technical support of the 

program; 

FSUE Scientific and Production 

Association “IMPULSE” (St. 

Petersburg) - development and 

upgrade of launch control and support 

equipment; 

 

State Enterprise Moscow Electrical and 

Mechanical Equipment Plant (Moscow) 

- modification of launch vehicle control 

system instrumentation; 

Ukraine 

 

 

National Space Agency of Ukraine 

(Kiev) - state support and supervision; 

State Design Bureau (SDB) Yuzhnoye 

(Dnepropetrovsk) - primary design and 

development organization for the 

launch vehicle and the entire Dnepr 

SLS; 

State Enterprise “Production 

Association Yuzniy Machine Building 

Plant” (Dnepropetrovsk) - primary 

manufacturing entity; 

 

Scientific and Production Enterprise 

KHARTRON-ARKOS (Kharkov) - 

primary entity for launch vehicle control 

system; 

Republic of Kazakhstan 

 

Aerospace Committee of the Ministry 

of Energy and Mineral Resources - 

state support and supervision; 

State Enterprise “INFRAKOS” 

(Baikonur) - participation in Dnepr 

Program activities at Baikonur 

Cosmodrome; 

 

State Enterprise “INFRAKOS-EKOS”, a 

subsidiary of “INFRAKOS” (Alma-Aty) – 

ecological support of the program. 

Additionally, ISC Kosmotras, on a contract 

basis, uses services of partners located 

outside of Russia and Ukraine for Dnepr 

Program marketing. 

Figure 2.2-1 shows the Dnepr Program 

operations structure. ISC Kosmotras 

principal office maintains contacts with 

launch customers, conducts preliminary 

evaluation of launch services provision 

opportunities, concludes contracts with 

customers, organizes their fulfillment by 

companies – subcontractors from Russia, 

Ukraine and Kazakhstan, works in 

cooperation with the Russian Ministry of 

Defense, Russian Aviation and Space 

Agency, National Space Agency of the 

Ukraine and National Aerospace 

Committee of the Republic of Kazakhstan, 

as well as with other governmental 

agencies of these countries. 

In August 1999 ISC Kosmotras 

established an ISC Kosmotras subsidiary 

that is located in the city of Kiev, Ukraine. 

It is licensed by the NSAU to conduct 

space related activities. Its primary 

mission is to participate in the process of 

establishing favorable conditions for 

promotion of Dnepr launch services to the 

market. 


13

Figure 2.2-1 Dnepr Program Operations Diagram 


14

3. Purpose, Composition and Principal 

Characteristics of Dnepr Space Launch 

System 

Dnepr SLS is designed for expedient, 

high-accuracy injection of various single or 

multiple spacecraft weighing up to 3.7 

metric tons into low-earth orbits inclined 

50.5, 64.5, 87.3 or 98 degrees. 

Dnepr SLS is developed on the basis of 

the SS-18 missile system, which is being 

decommissioned in the process of 

reduction and elimination of strategic 

offensive arms. In the course of the 

development of the system, the following 

major principles are being followed: 

maximum heritage with previously 

developed and proven systems and 

units; 

Dnepr LV is based on the SS-18 ICBM 

modified in order to ensure optimal 

spacecraft integration and injection with 

the minimum costs incurred. 

Launch complex is a combination of 

technologically and functionally 

interrelated systems, components, 

facilities and service lines required to 

maintain the readiness status of the 

launch vehicle, and to prepare and 

execute its launch. 

LV processing facility incorporates special 

structures and mobile processing 

equipment required to prepare the LV for 

launch. 

SC and SHM processing facility is 

designed to perform the following 

operations: 

 

utilization of proven work technologies; 

and 

 

acceptance, temporary storage and 

processing of the spacecraft; and 

utilization of LV standard flight 

trajectories and their associated drop 

zones. 

Dnepr SLS consists of the following 

elements: 

 

 

 

 

launch vehicle with the space head 

module; 

launch complex with the launch control 

center; 

launch vehicle, spacecraft and space 

head module processing facilities; and 

set of data collection and processing 

means. 

 

assembly and processing of the space 

head module consisting of the 

spacecraft, adapter, intermediate 

section, gas-dynamic shield and LV 

fairing. 

Available telemetry posts are used for 

receiving telemetry data during injection 

into orbit and for taking trajectory 

measurements. 

If the spacecraft is launched southward 

(87.3, 98 degrees inclination), telemetry 

posts located along the flight path on the 

territory of foreign countries may be used 

in order to ensure stable coverage. 

The sequence of ground telemetry post 

operation (including mobile telemetry 

assets) is contingent on the injection 

pattern of a specific spacecraft. 


15

Depending on the spacecraft launch 

program requirements 1 to 3 silo 

launchers can be used, which are capable 

of up to 25 launches per year (if the 

personnel work in one shift). 

The launch vehicle inside its launch 

canister, when fuelled and placed inside 

the silo launcher, can be on stand-by 

awaiting integration with the SHM and SC 

for an unlimited period of time throughout 

its operational life. 


16

4. Dnepr-1 Launch Vehicle 

4.1 General Description 

Dnepr-1 launch vehicle with a space head 

module is a basic modification of the 

liquid-fuelled SS-18 ICBM with a threestage-plus-SHM 

in-line configuration. 

General view of the launch vehicle with the 

space head module is shown in Figures 

4.1-1 and 4.1-2. 

The LV first and second stages are 

standard SS-18 elements and used 

without modification. 

First stage propulsion unit features four 

single-chamber motors, while the second 

stage propulsion unit is composed of a 

main single-chamber motor and a fourchamber 

thruster. 

The LV third stage is a modified standard 

SS-18 third stage equipped with a liquid 

propellant, two-mode propulsion unit that 

operates based on a “drag” scheme. 

Modifications involve only the control 

system in order to provide optimal flight 

software and electrical links with the 

spacecraft. 

SHM is attached to the third stage upper 

end. SHM consists of a spacecraft, 

intermediate section, adapter, either gasdynamic 

shield (GDS), or Encapsulated 

Payload Module (EPM), protective 

membrane and SS-18’s standard fairing. 

Dnepr LV features a standard inertial highprecision 

computer-based control system. 

The LV telemetry system ensures 

transmission of telemetric data from the 

LV up to SC separation (including the 

moment of separation) from the LV. 

The LV safety system ensures flight abort 

of the first and second stages in case of 

an emergency (loss of flight stability). This 

system is based on the safety system 

used for flight testing of the SS-18 ICBM. 

Dnepr LV is steam ejected from its 

transport and launch canister to a height of 

approximately 20 meters above the 

ground by means of activation of the black 

powder gas generator. The first stage 

propulsion unit is ignited upon the rocket 

ejection from the launch canister. 

Separation of stages and fairing follows 

the proven SS-18 procedures. Spacecraft 

separation from the third stage is done by 

the third stage taking away from the 

spacecraft by means of throttled-back 

operation of its motor. Prior to the 

spacecraft separation, the gas dynamic 

shield or EPM cover is jettisoned. 

Principal characteristics of Dnepr-1 LV 

with a single-tier space head module are 

given in Table 4.1-1. 

Dnepr-1 LV performance curves are 

presented in Figure 4.1.3. 


17

Table 4.1-1 Dnepr-1 Main Characteristics 

Liftoff mass (with the spacecraft mass of 2000 kg), kg 

1 st stage 208900 

2 nd stage 47380 

3 rd stage 6266 

Thrust in vacuum, tons 

1 st stage 461.2 

2 nd stage 77.5 

3 rd stage (primary mode/throttled 

back operation mode) 

1.9/0.8 

Propellant components for all stages 

Oxidizer 

Fuel 

Amyl 

Heptyl 

Effective propellant capacity, kg 

1 st stage 147900 

2 nd stage 36740 

3 rd stage 1910 

Flight reliability 0.97 

SC injection accuracy (Orbit altitude H cir =300 km) 

for orbit altitude , km ±4.0 

period of revolution, sec. + 3.0 

for inclination, degrees ±0.04 

for ascending node right ascension, 

degrees 

±0.05 

Orbit Inclination, degrees 50.5°; 64.5°; 

87.3°; 98° 

Figure 4.1-1 Dnepr-1 General View 

Figure 4.1-2 SS-18 1 st and 2 nd Stages 

Inside TLC 


18

G sc , kg 

4000 

3600 

3200 

2800 

2400 

2000 

i=50.5° 

i=64.5° 

i=87.3° 

i=98.0° 

1600 

1200 

800 

400 

0 

300 400 500 600 700 800 900 

Н cir , km 

Figure 4.1-3 Dnepr-1 Performance Curves for Circular Orbits 


19

Figure 4.1-4 Mission Profile of Dnepr-1 LV Carrying a Large Spacecraft 


20

Figure 4.1-5 Mission Profile of Dnepr-1 LV with a Group of Spacecraft 


21

4.2 Spacecraft Injection Accuracy 

In spacecraft mission profile to a circular 

orbit without application of yaw maneuver, 

the maximum deviations (with probability 

P=0.993) of orbital parameters from 

nominal values at the moment of 

spacecraft separation do not exceed the 

values given in Table 4.2-1 below. 

In case of multiple spacecraft injection, the 

orbit is calculated for each spacecraft, and 

the injection accuracy of a specific 

spacecraft may be specified. 

The data contained herein is of general 

character. For each specific flight, the 

composition and values of monitored 

parameters may be specified subject to 

launch mission and associated 

modifications of 3 rd stage control system 

operation pattern, required orbit 

parameters, payload mass (number of 

spacecraft being inserted), separated 

element drop zones, etc. 

4.3 Launch Vehicle Axes Definition 

The Launch Vehicle pitch and yaw planes 

as well as its rotation axes are defined by 

the LV center line as its longitudinal axis 

and by mutually perpendicular I-III and II- 

IV axes (see Figure 4.3.1): 

 

the pitch plane is defined by the LV 

longitudinal axis and I-III axis. 

I-III axis is positioned in such a way that 

the I-direction of the axis is pointing away 

from the launch site (lift-off point), whereas 

III-direction is pointing to the launch site; 

 

the yaw plane is defined by the LV 

longitudinal axis and II-IV axis. 

The LV angular displacements are 

determined (relative to an observer looking 

forward from the LV aft section) as follows: 

 

 

 

the LV makes pitch rotation around II- 

IV axis. The pitch positive movement 

displaces the LV nose section in the 

upward direction (towards direction III 

of I-III axis); 

the LV makes yaw rotation around I-III 

axis. The yaw positive movement 

displaces the LV nose section in the 

left-hand direction (towards direction II 

of II-IV axis); 

the LV makes roll rotation around the 

longitudinal axis. The roll positive 

movement turns the LV clockwise 

(from direction I of I-III axis to direction 

II of II-IV axis). 

Positive directions of the pitch, yaw and 

roll rotation are shown by arrows in Fig. 

4-3.1. 

Table 4.2-1 Spacecraft Injection Accuracy 

Orbital Parameter 

Typical circular orbits 

H=300 km, i=98 o H=600 km, i=98 o H=900 km, i=65 o 

Altitude, km ±4.0 ±5.5 ±10.0 

Period of revolution, sec. ±3.0 ±4.0 ±6.5 

Inclination, deg. ±0.040 ±0.045 ±0.050 

Right ascension of the ascending node, deg. ±0.050 ±0.060 ±0.070 


22

Positive Roll 

Direction 

Positive Pitch 

Direction 

Positive Yaw 

Direction 

Figure 4.3-1 Launch Vehicle Major Axes 


23

4.4 Space Head Module 

4.4.1 Space Head Module Design 

The spacecraft is installed inside the SHM. 

The SHM is composed of the fairing, 

cylindrical intermediate section, adapter, 

protective membrane, GDS or EPM. 

Fairing is a four-cone structure that has a 

longitudinal joint along stabilization axes I 

and III, which divides the fairing into two 

half-shells (sections) that are tied together 

by 28 pyro-devices. The fairing is installed 

atop the cylindrical intermediate section 

and attached to it by means of 8 pyrodevices. 

Upon the fairing separation 

command, the half-shell and 

fairing/intermediate section attachment 

pyros are activated, the half-shells are 

hinged by means of spring pushers 

installed at the bottom end of the fairing 

and, upon reaching a certain angle of turn, 

are separated from the intermediate 

section. 

Intermediate section is a cylindrical part 

that has a length of 2080 mm (standard 

size) and diameter of 3000 mm and 

incorporates two platforms – A (on upper 

extreme ring frame of which, the fairing is 

installed) and B (the bottom extreme ring 

frame of which is attached to the 3 rd 

stage). Both platforms are interconnected 

by 6 pyro-devices. The length of 

intermediate section can be extended by 

850 – 2000 mm to incorporate a bigger 

spacecraft. 

Adapter is a newly developed conical 

structure. Attached to the adapter upper 

flange are the spacecraft and a special 

protective membrane that isolates the 

payload envelope from the control and 

telemetry system instrumentation 

compartment located under the adapter. 

The adapter bottom flange sits on the 

lower extreme ring frame of the platform B. 

In necessary, an additional adapter (or 

several adapters) that is already integrated 

with the spacecraft, may be placed on the 

standard adapter. 

SHM is available in two configurations: 

Configuration 1 – a newly developed 

intermediate section with GDS; 

Configuration 2 – an intermediate section 

consisting of standard platforms A and B 

plus EPM. 

Both configurations of SHM use the 

standard fairing and a newly developed 

adapter. 

For SC protection against the 3 rd stage 

motor plume, the Configuration 1 utilizes 

the GDS that is attached to the Platform A 

upper ring frame and separated prior to 

the SC release. 

For the above purpose, the Configuration 

2 uses the EPM, the cover of which is 

separated prior to the spacecraft release, 

i.e. similarly to the GDS. 

Layout schematics of the standard length 

SHM (both with GDS, and EPM) is shown 

in Figure 4.4.1.1 and 4.4.1.2 respectively. 

Layout schematic of SHM Configuration 1, 

the length of which is extended by 850 

mm, is shown in Figure 4.4.1.3. 

SHM design allows for multi-tier spacecraft 

layout. One of the options for such layout 

is shown in Figure 4.4.1.4. 


24

Figure 4.4.1-1 SHM Configuration 1 – 

Standard Length 


Extended by 850 mm 


Standard Length 

Figure 4.4.1-4 SHM with 2-tier Layout 


25

4.4.2 Payload Envelope 

Payload envelope is a volume within the 

SHM, which is designed for 

accommodation of spacecraft. 

Spacecraft dimensions (including all of its 

protruding elements) must fit within the 

specified payload envelope, given all 

possible deviations and displacements 

from the nominal position during the 

ground testing and in flight. 

The size of payload envelope within the 

standard SHM (length – 5250 mm) and 

adapter (H = 550 mm) is shown in Figures 

4.4.2.1 and 4.4.2.2 (i.e. for SHM 

Configuration 1 with GDS and SHM 

Configuration 2 with EPM respectively). 

Figure 4.4.2.3 shows the size of payload 

envelope within the SHM extended by 850 

mm (maximum possible extension of the 

SHM length is 2000 mm). 

When designing the interface between the 

launch vehicle and a specific spacecraft, 

the size and configuration of the payload 

envelope may be specified. 


26

Figure 4.4.2-1 Payload Envelope Available within SHM Configuration 1 with Standard 

Adapter and GDS 


27

Figure 4.4.2-2 Payload Envelope Available within SHM Configuration 2 with Standard 

Adapter and EPM 


28

Figure 4.4.2-3 Payload Envelope Available within SHM Configuration 1 Extended by 

850 mm with Standard Adapter and GDS 


29

4.5 Launch Vehicle Flight Reliability 

SS-18 in its basic configuration has been 

in operation since mid-70s. Throughout 

the entire period of its operation, a number 

of measures have been taken to maintain 

the technical condition of the rockets, 

launch and processing equipment. These 

measures include the following: 

 

 

 

 

Regular inspections of the operability 

of the systems, units and equipment; 

Scheduled (routine) maintenance; 

Replacement of faulty units, modules 

and components; and 

Periodical checks of the readiness of 

all systems, units and equipment for 

the performance of their basic 

functions. 

The effectiveness and adequacy of the 

above measures have been demonstrated 

by the successful launches of the rockets 

with the different storage periods at 

various times during the SS-18 operational 

life. 

In order to monitor the actual degree of the 

launch vehicle reliability, the leading 

institutes of the Russian Ministry of 

Defense and Russian Aviation and Space 

Agency, as well as State Design Bureau 

Yuzhnoye as the principal developer of the 

system, conduct annual independent 

evaluation of the reliability data values, of 

their compliance with the established 

requirements and also develop programs 

of work necessary to ensure the required 

level of reliability. The implementation of 

this work ensures and allows to maintain 

the high level of the rocket reliability 

throughout the entire period of its 

operation. Currently, the mission reliability 

factor for the rocket is 0.97, which has 

been established through a great number 

of successful launches. 

The total of 159 launches of the SS-18 

were carried out as of October 2001 with 

only 4 of the launches encountering 

malfunctions or anomalies of certain 

systems. The causes of the failures have 

been unequivocally established. They 

were due to isolated manufacturing 

defects. Based on the analysis of these 

defects, certain measures were taken to 

enhance rocket fabrication quality and 

ensure full compliance with engineering 

documentation during fabrication. The 

effectiveness of the measures taken was 

proved through subsequent successful 

launches. No similar recurring 

malfunctions were encountered. 8 SS-18 

launches have been conducted from 1990 

to 2000, all of them successfully 

accomplished their mission. 

The conversion of the SS-18 into Dnepr-1 

launch vehicle requires minor design 

modifications associated with the 

spacecraft integration with the launch 

vehicle and spacecraft separation process. 

The modified components undergo 

vigorous ground testing that enables to 

maintain the achieved mission reliability 

level. The above-mentioned 8 launches 

included two launches in Dnepr-1 

configuration that were performed in 1999 

– 2000. 


30

5. Baikonur Cosmodrome 

Dnepr SLS operates from Baikonur 

Cosmodrome, one of the largest 

spaceports in the world. 

Baikonur Cosmodrome is located in 

Kazakhstan east of the Aral Sea (63 E and 

46 N) in the semiarid zone with sharply 

continental climate. Typical for this area 

are hot dry summers and frosty winters 

with strong winds and little precipitation. In 

summer, the temperature can rise up to 

plus 45 0 C and in winter, it can drop to 

minus 40 0 C. Yearly average temperature 

is approximately plus 13 0 C. 

Figure 5-2 Syrdarya River 

The Cosmodrome is located sufficiently far 

away from the large centers of population. 

This fact ensures safety of the launches 

and facilitates the task of creating buffer 

areas, which are used as launch vehicle 

stage drop zones. Another advantage of 

the Cosmodrome location is the fact that 

typical for this location is a great deal of 

clear days on a yearly basis. 

The residential area of the Cosmodrome is 

the town of Baikonur located on the right 

bank of the Syrdarya River. Directly 

adjacent to the town are the village of 

Tyuratam and the railway station of the 

same name. Two Baikonur airports are 

connected with the city of Moscow by 

regular flights. 

Figure 5-3 Cosmodrome Residential Area 

– Town of Baikonur 

These airports are equipped to 

accommodate any types of cargo and 

passenger planes. 

Figure 5-1 Yubileiniy Airfield 

The town has the following facilities: a 

movie theatre, a stadium, a swimming 


31

pool, TV center, medical facilities, cafes 

and restaurants as well as international 

telephone, telex and facsimile services. 

Baikonur visitors can stay at several hotels 

that offer single and double rooms. 

to the above-mentioned facilities, have a 

bed room and a private office. The cost of 

accommodation is $30 – 80 per day at 

Baikonur and Kosmonavt hotels and $220 

– 250 per day at Sputnik hotel. 

Accommodations are offered for 

spacecraft processing specialists and 

launch guests at Sputnik, Baikonur and 

Kosmonavt hotels located in the northern 

part of the town. 

All Russian and foreign cosmonauts stay 

at Kosmonavt hotel during preparation for 

space flight. 

The rooms have air conditioning, a 

television set, a refrigerator and a bath 

room. Deluxe accommodations, in addition 

Figure 5-4 Kosmonavt Hotel 

Figure 5-5 Baikonur Hotel 

Figure 5-6 Facilities of Sputnik Hotel 


32

ATMs and exchange offices are available 

in Baikonur for currency exchange and 

cash withdrawal. Both Russian and 

Kazakhstan currencies are accepted. Most 

payments are made by cash. 

The Cosmodrome is connected with other 

cities and countries by air, rail and road 

transportation. There is an extensive 

network of motor and railroads on the 

territory of the Cosmodrome. 

The town provides centralized water (hot 

and cold) and 220V/50Hz power supply. 

The basic accommodation for the 

customer’s personnel is assumed to be 

Kosmonavt hotel. 

Table 5.1 shows typical routes of 

transportation at Baikonur Cosmodrome. 

Table 5-1 Typical Routes of Transportation 

at Baikonur Cosmodrome 

Krainiy Airport Kosmonavt Hotel 6 km 

Yubileiniy Airport Kosmonavt Hotel 55 km 

Tyuratam Railway 

Station 

Kosmonavt Hotel 

3 km 

Kosmonavt Hotel Launch Site 45 km 

Kosmonavt Hotel Facility 31, 

SC Fuelling Facility 

60 km 

Kosmonavt Hotel Facility 42, AITB 55 km 

Minivans and sedan cars are available for 

transportation of personnel. The presence 

of local escorts to accompany transport 

vehicles is a must. 

Accommodation for VIPs is available at 5- 

star Sputnik hotel. Police escorts can be 

arranged for VIP transportation. 

During spacecraft preparation for launch at 

S/C processing facility and launch 

complex, the Customer can be provided 

with the following telecommunication 

services: 

local/international telephone/facsimile 

communication; 

internal technological and public 

address system between work sites at 

processing facility premises (intercom); 

mobile radio and radiotelephone 

communication; 

 

 

communication channels to transmit 

UTC and pre-launch countdown 

signals; 

access to the Internet, etc. 

Dnepr SLS telecommunication system is 

integrated into the single distributed 

information network of the Baikonur 

Cosmodrome based on the ISDN 

principles by wire cable and radio relay 

communication channels network. This 

network covers the premises and facilities 

of the launch complex (Sites 103, 106, 

109, 111/2), S/C processing facility (Sites 

31, 42) and other Cosmodrome facilities. 

International communications are provided 

by a satellite segment based on the 

Intelsat and Inmarsat type ground stations 

with direct access to the Moscow 

telephone network. 

The mobile communication is provided by 

the trunking radio stations covering the 

complete Cosmodrome area, the GSM- 

900 cellular communication is provided at 

the most of the cosmodrome and adjacent 

town areas. 


33

A fully equipped military hospital is located 

near the Kosmonavt hotel. Its personnel 

has the required qualification and 

experience. First medical aid is available 

at work premises and sites. 

Upon customer’s request, medical 

personnel and equipment can be placed 

on stand-by at spacecraft processing 

facility and in emergency circumstances, 

the patient may be evacuated to a 

European medical facility. 


34

6. Dnepr SLS Ground Infrastructure 

6.1 Elements and General Diagram of 

Ground Infrastructure 

The Dnepr SLS ground infrastructure 

includes the following facilities: 

 

 

 

 

SC and SHM Processing Facilities; 

SC Fuelling Station; 

Launch complex; 

LV Processing Facility; 

Figure 6.1-1 shows infrastructure facilities 

at Baikonur Cosmodrome, which are used 

by ISC Kosmotras for Dnepr launch 

campaigns. 

6.2 SC and SHM Processing Facility 

SC and SHM processing facility is 

designed to perform the following 

operations: 

Receiving, temporary storage and 

processing of the spacecraft, including, 

if necessary, its filling with compressed 

gases and propellant components; and 

 

Fueling station for storage, preparation 

and discharge of rocket propellant 

components; 

 

Assembly and processing of the space 

head module (consisting of spacecraft, 

adapter and launch vehicle fairing). 

97 90 

93 

94 

131 

91 

92 

82 

95 

182 

81 

175 

200 

251 

Yubileiny Airfield 

250 

Dnepr LV 

Silo Launchers 

SHM and SC 

Processing Facility 

Railways 

Freeway 

Motor roads 

31 32 

104 

110 

43 41 

118 

119 

112A 

112 

254 51 

2 1 

113 

18 

111/2 

103 

106, 109 

Yuzhnaya 

Hotel 

45 

42 44 

W 

N 

E 

5 

Dnepr LV Control and 

Observation Post 

S 

Dyurmentyube 

3K 

21 

3G 

23 

3R 

Kosmonavt, 

Baikonur, 

Sputnik 

Hotels 

Krayniy 

Airfield 

Tyuratam 

15 

10 

BAIKONUR 

Figure 6.1-1 Dnepr SLS at Baikonur Cosmodrome 


35

Containers with the spacecraft and 

spacecraft ground processing equipment 

(GPE) can arrive Baikonur via two airports 

- Yubileiniy and Krainiy as well as via 

TyuraTam railway station. 

If a spacecraft needs to be kept in certain 

temperature conditions, it will arrive at 

Yubileiniy airport, from where it will be 

transported to SC and SHM processing 

facilities by special environmentally 

controlled railcars for subsequent 

operations. 

If there is no strict temperature 

requirement, the spacecraft and SC GPE 

will be transported by road vehicles or 

general purpose railcars. 

Two airports (Yubileiniy and Krainiy) are 

capable of receiving all passenger and 

cargo airplanes, including heavy ones (An- 

124 and Boeing 747) and possess all 

necessary ground support equipment for 

handling operations. 

Directly adjacent to Yubileiniy airport is a 

rail line that can be used for transportation 

of spacecraft and SC GPE to SC and SHM 

processing facility. 

Depending on spacecraft processing 

requirements, its weight and dimensional 

characteristics and the cost of operation, 

ISC Kosmotras offers customers 

assembly, integration and test buildings 

(AITB) for spacecraft processing at Site 42 

and/or Site 31. 

If necessary, other AITBs available at 

Baikonur Cosmodrome may be used for 

spacecraft processing. 

Premises and equipment located inside 

Figure 6.2-1 Layout of AITB, Site 42 


36

the Zenith LV AITB were used for 

spacecraft processing in the course of first 

two Dnepr launch campaigns (Site 42). A 

layout of AITB, Site 42 is shown in Figure 

6.2-1. Technical data for clean room 

facility (primary place for spacecraft 

processing and assembly of EPM) and 

other AITB premises is given below. 

Clean room facility located inside the AITB 

is divided into two clean rooms with the 

aggregate floor area of 212 sq. meters and 

an airlock compartment of 32.6 sq. meters. 

Telphers with lifting capacity of 1000 kgf 

are available in the airlock compartment 

and the two clean rooms. Maximum height 

to the telpher hook is 8.2 meters. The 

sizes of the airlock compartment gate, the 

gate between clean rooms and the gate 

opposite the airlock compartment are 

5.2x6.3 m, 5.2x8.2 m and 5.2x6.3 m 

respectively. The temperature inside the 

clean room facility is maintained within the 

range of 21.1°C - 26.7°C with the relative 

humidity being 30-60%. Cleanliness class 

is 100,000 (US FED-STD-209E). Clean 

room facility has an anti-static floor, a lowimpedance 

grounding contour (up to 4 

ohms) and 12V/220V AC 50 Hz electric 

outlets. Illumination of airlock compartment 

is 300 lux and of clean rooms is 500 lux. 

Spacecraft checkout equipment (COE) 

room has the aggregate floor area of 53 

sq. meters (29 sq. meters and 24 sq. 

meters). 

The AITB has office premises for 

personnel involved in spacecraft 

processing and assembly (4 rooms, 21 m 2 

each), which are equipped with required 

communications means. 

The AITB is equipped with the following 

utilities: uninterrupted power supply (with 

the same characteristics as both in Russia 

and the US), heating, air conditioning and 

ventilation, water supply, sewer, security 

and fire alarms, and various 

communication systems. 

Near the clean room facility, there is a 

Figure 6.2-2 Layout of AITB, Site 31 


37

place where the SHM or EPM is 

assembled and loaded into the transporter 

erector. 

If the weight and dimensions of the of 

spacecraft do not allow to use the AITB at 

site 42 or the spacecraft filling with 

propellant is required during its 

processing, ISC Kosmotras offers to use 

the AITB and fuelling station located at 

Site 32, shown in Figure 6.2-2. 

This AITB consists of three areas A, B and 

C. 

 

 

 

Area A is designed for integration of 

EPM and for electrical checks of 

spacecraft as part of the EPM. To 

maintain the required cleanliness 

conditions during operations, the Area 

A is equipped with “air shower”; 

Area B is designed for the spacecraft 

processing and assembly of SHM; 

Area C is designed for receiving and 

temporary storage of shipping 

containers with SC and SC GPE, of 

SHM platforms (SC adapters) and 

launch vehicle fairing, as well as for 

handling operations including the 

loading of the integrated SHM into the 

transporter erector for transportation to 

the silo launcher. 

Main characteristics of areas A, B and C 

are given in Table 6.2-1 below. 

Facility 40D within the Area B consists of 

hall 119, 119A and 119B, the area of 

which is 240, 100 and 100 m 2 respectively. 

Hall 119 will incorporate the spacecraft, 

hall 119A will host spacecraft COE, and 

hall 119B will be for mechanical 

equipment, devices and consumables. 

Hall 119 is equipped with compressed gas 

filling system, including filling with 

Nitrogen. This Nitrogen will be supplied 

under the pressure of 40 MPa (standard 

ОСТ 92-1577-78) and have the following 

parameters: 

nitrogen content – no less than 

99.95%; 

oxygen content – no more than 0.05%; 

water vapor – 0.004%. 

Table 6.2-1 Characteristics of Area A, B and C, AITB at Site 31 

Dimensions 

Cleanliness Class (US 

standard 209E) 

Area A B C 

Length 56.0 m 

Width 11.5 m 

100,000, 

if necessary 30,000 

Characteristics 

Length 56.0 m 

Width 18.5 m 

100,000, 

if necessary 30,000 

Height to Crane Hook 13.8 m 13.8 м 13.8 м 

Length 63.0 m 

Width 30.0 m 

Crane Lifting Capacity 10 and 50 tons 10 and 50 tons 10 and 50 tons 

Temperature 18-25°C 18-25°C 18-25°C 

Relative Humidity at 20 °C 30-60% 30-60% - 

- 


38

Hall 119 has the following power supply 

system: 

Fuelling station is equipped with following 

systems: 

 

120 V 20 A 50 Hz; 

 

oxidizer filling system; 

 

208 V 30 A 50 Hz; 

 

fuel filling system; 

 

15 kWt 50 Hz. 

 

fuelling remote control system; 

Hall 119 is equipped with video cameras. 

The personnel directly involved in 

spacecraft processing operations at Halls 

119, 119A and 119B enter area B through 

air showers located at both ends of the 

AITB and separating “clean” and “not 

clean” areas. The personnel not 

completely involved in operations with the 

spacecraft, who operate communications 

means or computers are located in office 

premises of the “not clean” area. To 

support operations with the spacecraft in 

“clean” area, the office premises are 

connected with hall 119 by the required 

number of cables run through hermetically 

sealed inlets. 

Work place for the SHM integration, 

processing and checks is also set up in 

area B. 

The AITB is equipped with the following 

utilities: uninterrupted power supply (with 

the same characteristics as both in Russia 

and the US), heating, air conditioning and 

ventilation, water supply, sewer, security 

and fire alarms, and various 

communication systems. 

6.3 Spacecraft Fuelling Station 

SС fueling station is located in close 

vicinity to the AITB and is used for filling 

spacecraft with liquid propellant and 

gases. 

 

system for collection and incineration 

of propellant vapors and sewage; 

fuelling equipment neutralization 

system; 

temperature and humidity control 

system; 

 

 

 

 

vacuumization system; 

gas control system; 

fire fighting system; 

set of scales. 

Depending of the SC fuelling 

requirements, the propellant components 

undergo necessary processing: 

 

 

 

 

filtration through 20 or 5 micron filters; 

providing required temperature and 

humidity of propellant; 

saturation with nitrogen or helium; and 

degassing. 

The fueling station is a heated 97-m long 

and 41-m wide building. The building is 

divided into three sections: section No. 1 – 

filling with fuel, section No. 2 - filling with 

compressed gas, and section No. 3 – 

filling with oxidizer. 


39

The external doorways and openings 

between the sections have hermetic seals 

– rolling hermetic doors 5.5 meters long 

and 7 meters high. The sections are 

equipped with 15-ton capacity bridge 

cranes. For the purposes of transportation, 

there is a rail line running through the 

filling rooms. The temperature in the filling 

rooms is kept at +5 0 C to +35 0 C. Filling 

rooms are equipped with “clean tents”, 

which provide the cleanliness class 

100,000 in accordance with US Federal 

Standard 209 A. 

6.4 Launch Complex 

Launch complex is a combination of 

service facilities, systems and lines, which 

ensure accomplishing of the following 

tasks: 

 

 

 

Achieving launch readiness status for 

the Dnepr LV and SC; 

Continuous and periodical automated 

remote control of LV, SC and silo 

launcher equipment parameters; and 

LV launch. 

The launch complex includes: 

 

 

silo launchers; 

launch control center (LCC); 

Figure 6.4-1 Dnepr Launch Complex 

Facilities (Sites 106 and 109) 

 

roads and technical facilities available 

between sites. 

If active electrical interface is used or the 

Customer needs to monitor the SC status 

up to launch, the underground Facility 25 

can be used for accommodation of the SC 

COE. This facility is located in the close 

vicinity of the silo launcher and is 

connected with it by a 80 m utility tunnel. 

Facility 25 is a reinforced concrete 

structure consisting of a number of rooms. 

In particular, a 67 m 2 room (5.7 by 11.8 m 

and 2.4 m high) will be available for the 

Customer’s COE. 

Facility 25 has a 220/380V 3-phase 15 

kWt power supply system. Required 

temperature (+ 15 - +25 0 C) is maintained 

by electric heaters. 

 

 

standard internal power supply system; 

communication and control cable lines 

running between sites; 

Facility 25 can be equipped with various 

communications means to communicate 

with the LCC and other services both at 

Baikonur Cosmodrome and outside of it, 

as required by the Customer. ISDN line 

(64kbit/sec.) can be made available for 

Facility 25, if necessary. 


40

In addition to that, the Customer may be 

offered a so-called customer monitoring 

post, located near the LCC at a distance of 

about 7 km from the launch site and 

Facility 25. These premises may 

accommodate customer’s personnel, who 

monitor the SC parameters up to the 

launch. If the parameters being monitored 

differ from the nominal values, the 

Customer will have an opportunity to 

suspend the pre-launch sequence no later 

than 3 minutes prior to launch in order to 

save the spacecraft. 

Figure 6.4-2 Transporter Erector 

A transporter erector, equipped with a 

system that allows to maintain certain 

temperature and humidity conditions, is 

used for transportation of the SHM 

containing the SC from AITB to the silo 

launcher for integration with the launch 

vehicle. 

The launch of the LV and control over the 

launch command execution are carried out 

via wire communications by the remote 

control system, the equipment of which is 

located at the LCC. 

If the launch is cancelled or postponed, or 

it is necessary to detach the SHM from the 

launch vehicle, the operations are 

conducted in a sequence reverse to the 

SHM/LV integration sequence. 

In case of spacecraft on-board equipment 

malfunction, or if it is necessary to recheck 

the spacecraft at the SC processing 

facility, operations of it’s delivery to the 

processing facility (Site 42 or 31) are 

performed in a reverse sequence of SC 

integration with the SHM. 


41

7. Baikonur Operations Flow 

 

SC / SHM integration; 

Spacecraft and launch vehicle preparation 

operations are divided into several stages, 

which may overlap depending on the 

spacecraft preparation requirements: 

 

 

 

preparation of LV for integration with 

SHM; 

SC processing; 

preparation of SHM for mating with the 

SC; 

 

 

 

transportation of SHM containing SC 

by transporter erector to the silo 

launcher; 

SHM/LV integration and LV preparation 

for launch; 

pre-launch operations and LV launch. 

Figure 7-1 shows work schedule for 

preparation and launch of a SC by 

Dnepr-1 LV (all dates are tentative). 

# Stage 

Duration, days 

1 

Spacecraft processing 

≥20 

2 

Preparation of SHM for 

mating with the SC 

3 SC/SHM integration and 

transportation to launch silo 

4 

Preparation of LV for 

integration with SHM 

22 

9 

7 

5 SHM/LV integration 

1 

6 

Pre-launch operations and 

LV launch 

6 

Figure 7-1 Spacecraft Processing and Launch Schedule 


42

7.1 Preparation of LV for Integration with 

SHM 

LV preparation for integration with the 

SHM is done as follows: 

installation of the launch vehicle 

transport and launch canister 

containing the 1 st and 2 nd stages into 

the silo launcher; 

equipping 3 rd stage with telemetry 

system; 

 

 

 

 

filling of 3 rd stage with propellant; 

mating 3 rd stage and LV inside the silo; 

set-up of work place for SC checks in 

Facility 25 (if active electrical interface 

is used); 

checkout of the launch vehicle with the 

electrical equivalent of the SHM, check 

of SC/work place-Facility 25 line; and 

fueling of stages 1 and 2 with 

propellant components. 

Duration of LV preparation for integration 

with SHM is about 21 days. The fueled LV 

can await integration with the SHM inside 

the silo for an unlimited period of time 

(within its operational life). 

7.2 SC Processing 

Containers with the spacecraft and its 

ground processing equipment (GPE) 

arrive Baikonur by air or by rail, in 

accordance with the launch campaign 

schedule agreed between the Customer 

and ISC Kosmotras, normally, no less than 

20 days prior to launch. 

Container off-loading is done by lifting 

gear of the airplane, by road cranes or 

forklifts using pull ropes, cross-arms, 

pallets and other Customer’s devices. 

Arrived containers with the spacecraft and 

GPE will be delivered to the SC/SHM 

processing facility for subsequent 

operations by road or rail, on common 

carrier rail flatcars in shrouded condition to 

protect from direct impact of precipitation. 

Figure 7.1-1 Loading of TLC Containing 

LV 1 st and 2 nd Stages into Silo Launcher 

Figure 7.2-1 Off-loading of Spacecraft 

Container from the Plane 


43

If necessary, the containers can be 

transported from Yubileiniy airport on 

special environmentally controlled railcars. 

At customer’s request, the ground 

transport loads (g-loads, accelerations, 

temperature) can be controlled during 

transportation and the results made 

available to the customer. 

At SC and SHM processing facility, the 

containers with the spacecraft and GPE 

are off-loaded inside the AITB. GPE is 

deployed at designated work places and 

the spacecraft is brought into the clean 

room (clean room inside AITB at Site 42 or 

Area B inside AITB at Site 31). 

Spacecraft processing is done based on 

the spacecraft authority requirements, 

including its filling with propellant 

components and gases at the fuelling 

station, if necessary. To ensure the 

required temperature environment during 

the spacecraft transportation from the 

AITB to the fuelling station, the spacecraft 

is delivered to the fuelling station inside its 

container on a special environmentally 

controlled railcar. 

The customer may use its own propellant 

(hydrazine). 

7.3 Preparation of SHM for Mating with 

SC 

Upon completion of the spacecraft 

processing (if processing is done at Site 

42), the SHM or EPM is delivered to the 

AITB by a special road vehicle. 

Upon delivery, the SHM is off-loaded and 

the fairing is detached. 

The SHM or EPM is brought inside the 

clean room by a special rail trolley. 

Adapter(s) is prepared for mating with the 

spacecraft. Spacecraft adapter(s) may be 

detached from SHM or EPM and placed 

on special supports. 

7.4 SC / SHM Integration 

Figure 7.2-2 Spacecraft Processing 

Counterparts of separation mechanisms 

are mounted on the spacecraft. 


44

Spacecraft are mounted on adapters 

manually or by telphers available at clean 

room; spacecraft electrical checks and 

battery charging (if necessary) are 

conducted. Then the spacecraft with their 

adapters attached are mated with the 

SHM or EPM. 

Final assembly of SHM or EPM is carried 

out. SHM or EPM is then taken out of the 

clean room by a rail trolley. Fairing is 

attached to the SHM. 

SHM or EPM is loaded into a special road 

vehicle and transported to the AITB at Site 

31 for SHM final assembly, its electrical 

checks and preparation for integration with 

the launch vehicle. 

If the spacecraft is processed at Site 31, 

the assembly of the SHM (with or without 

EPM) is conducted at the Area B of the 

AITB at Site 31 following the same 

sequence as for the Site 42, except for the 

transportation of the SHM (EPM) between 

sites. 

7.5 Transportation of SHM Containing 

SC by Transporter Erector to the Silo 

Launcher 

Assembled and processed SHM is loaded 

into the transporter-erector and delivered 

to the launch complex. During the 

transportation, the required temperature 

and humidity conditions are maintained 

around the spacecraft by means of 

transporter erector climate control system. 

SHM delivery time to the launch complex 

does not exceed 3 hours. 

Figure 7.4-1 SHM Integration 

Figure 7.5-1 SHM Loading into 

Transporter Erector 


45

7.6 SHM/LV Integration and LV 

Preparation for Launch 

Upon arrival at the launch complex, the 

SHM, by means of transporter erector, is 

mated with the LV installed inside the silo. 

SHM/LV mating time does not exceed 3 

hours. 

Electrical checks of the LV with SHM, data 

analysis, final launch preparation 

operations are performed. Time period 

from the completion of SHM/LV mating to 

launch does not exceed 6 days. 

If necessary, during LV preparation for 

launch, continuous (periodical) monitoring 

of the spacecraft from Facility 25 can be 

organized. Monitoring stops 30-40 minutes 

prior to launch since all personnel must be 

evacuated from danger zone to the vicinity 

of the LCC. 24-hour stand-by duty can be 

organized at the launch complex to 

support operations with the spacecraft 

COE. 

Figure 7.6-1 SHM/LV Integration Inside 

Silo Launcher 


46

Transporter Erector 

Silo Launcher 

Space Head 

Module 

Figure 7.6-2 SHM Loading into Silo Launcher 

7.7 Pre-launch Operations and LV 

Launch 

The LV launch is conducted from the LCC 

(Site 111/2). The launch command is 

issued at the pre-determined moment of 

time, and the following operations are 

performed prior to the launch command 

issuance: 

 

 

final operations are conducted at the 

LCC (3 hours); 

2 hours prior to launch, the preparation 

of the telemetry system ground 

equipment begins; 


47

1.5 hours prior to launch the silo door 

is opened and mobile equipment in the 

vicinity of silo is evacuated; 

1 hour prior to launch, the on-board 

telemetry system is activated (for about 

10 minutes) and the telemetry data 

receipt by the ground stations is 

verified; 

20 minutes prior to launch, the onboard 

telemetry system is activated 

again and the personnel are 

evacuated; 

3 minutes before launch by a 

command from the launch control post, 

the on-board telemetry system starts 

getting power from the LV on-board 

power supply system, the ground 

stations start receiving, recording and 

processing the actual pre-launch and 

flight telemetry; 

 

LV is launched. 

After the launch, the launch crew and 

telemetry system crew return to the silo 

launcher and conduct final operations with 

the telemetry system checkout equipment 

and examination of the silo launcher. 

Figure 7.7-1 Dnepr LV Liftoff 


48

8. SC/LV Interfaces 

8.1 Mechanical Interface 

Spacecraft is attached to the launch 

vehicle adapter by means of pyro-devices. 

The number and type of pyro-devices is 

contingent on a number of parameters 

(spacecraft weight, the diameter on which 

the attachment points are located, etc.) 

and is agreed upon with the customer for 

the specific spacecraft. 

For some types of spacecraft, the 

adapters that were used in previous 

launches, can be applied. They include 

adapters for 10-20 kg spacecraft, 50-70 kg 

spacecraft and 300-400 kg spacecraft. 

These adapters can serve as prototypes 

for the development of new ones. 

Typical design of the spacecraft/standard 

adapter attachment point is shown in 

Figure 8.1-1. 

To verify the spacecraft separation from 

the Space Head Module, separation 

switches mounted on the adapter are 

used. 

To disconnect electrical connectors that 

ensure the LV/SC electrical links, 

separation mechanisms are used, which 

are activated prior to the operation of the 

separation system. 

Pyro-devices, separation switches and 

mechanisms used, have undergone all 

necessary ground and flight testing and 

are highly reliable. 

Separation system activation equipment 

and cables are installed on the Space 

Head Module and adapter in compliance 

with all operational requirements, which 

precludes any damage to or collision with 

the elements being separated and 

satisfies high reliability requirements of 

electrical command relay that was 

confirmed by ground and flight testing. 

If necessary, and subject to mutual 

agreement, the adapter can be supplied to 

the spacecraft authority by ISC Kosmotras 

and SDB Yuzhnoye for fit-check testing. 

It should be noted that the separation 

systems used for SC/Dnepr-1 LV 

separation have no spring pushers, since 

the spacecraft separation is done by 

taking away the upper stage from the 

spacecraft by means of the upper stage 

motor throttled-back operation, which 

ensures minimum disturbances on the 

spacecraft during separation. 

Angular stabilization errors of the launch 

vehicle 3 rd stage with the Space Head 

Module at the moment of issuing the 

spacecraft separation command, do not 

exceed: 

 

 

+ 1.5 degrees for angles of pitch, yaw 

and roll; and 

+ 0.5 degrees per second for rates of 

pitch, yaw and roll. 

Spacecraft disturbances due to process of 

separation are dependant on the inertial 

characteristics of the spacecraft and the 

3 rd stage with the Space Head Module 

(including the spacecraft), on the type, 

number, location and characteristics of the 

spacecraft/space head module attachment 

joints. Spacecraft (with the weight 

exceeding 300 kg) angular rates after 

separation due to stabilization errors and 

disturbances induced by the separation 

process are as follows: 

ω x ‹ 2.0 degrees per second; 

ω y ‹ 3.0 degrees per second; 

ω z ‹ 3.0 degrees per second. 


49

Analysis of the relative motion of the 

spacecraft being separated and SHM 

elements confirms the spacecraft 

separation safety, i.e. indicates no collision 

among the spacecraft themselves and with 

the SHM elements. 

* - 20 by 20 mm seats, where separation switches will rest, should be available on the spacecraft. 

Figure 8.1-1 Typical Design of Spacecraft/Standard Adapter Attachment Point 


50

8.2 Electrical Interface 

Two types of SC/LV electrical interface are 

possible when launching a spacecraft on 

Dnepr-1 LV: 

Passive electrical interface – no 

electrical connections between LV and 

SC. No power is supplied to the 

spacecraft during ascent. Upon 

spacecraft separation, its separation 

switch is activated and its power supply 

system turns on. The LV has 

spacecraft separation sensors which, 

upon separation, generate a telemetry 

signal that is transmitted to the ground; 

and 

 

Active electrical interface – electrical 

connections available between SC and 

LV that are used prior to (SC-SHM-LV- 

TLC-SC COE) and during launch (SC- 

SHM-LV) as well as during injection 

into orbit (SC-SHM). 

The type of the electrical interface is 

agreed upon between the Customer and 

ISC Kosmotras. The above interface types 

differ by their complexity and cost of their 

technical realization. 

SC/LV electrical links ensure the following: 

 

 

transmission, if necessary, of telemetry 

data on the status of the SC on-board 

systems during ascent as well as 

during electrical checks of the SC 

integrated with the LV inside the silo 

launcher; 

issuance of commands by the LV 

control system to the SC on-board 

instrumentation during ascent as well 

as during electrical checks of the SC 

integrated with the LV inside the silo 

launcher; 

supply of power for SC battery 

charging, maintaining required 

temperature and humidity conditions of 

the SC bus and for the SC command 

and control interface, prior to launch 

when the SC is integrated with the LV 

inside the silo launcher. 

Existing through circuits of the LV control 

system and TLC that can maintain about 

20 various transmission lines are used for 

LV/SC and LV/SC checkout equipment 

electrical links. Diagram of 

Spacecraft/Spacecraft ground equipment 

links is shown in Figure 8.2-1. 

Through circuits run via umbilical 

connectors on the LV frame, which are 

ruptured prior to launch – 3.5 seconds 

before LV starts moving inside the TLC. 

Spacecraft ground equipment is 

connected to the LV and TLC through 

circuits by means of additional connectors 

mounted on the TLC and cables running 

from the silo to Facility 25. 

A 32-pin electrical connector (ГШР8) is 

available at LV/SHM interface, the pins of 

which have 0.75 mm 2 wires running from a 

3-pin electrical connector (ШР8) that is 

installed at ground/LV side connector 

board and allows to pass the total 

amperage of up to 50 A. On the SHM, 

between the ГШР8 connector and two 

spacecraft umbilical connectors, a new 

cable will be run for battery charging and 

maintaining the spacecraft on-board 

equipment required temperature and 

humidity conditions. 

In addition to that, LV/SHM interface has 

three more 102-pin electrical connectors 

(ГШР1, ГШР2, ГШР3), some pins of which 

have 0.2 mm 2 and 0.35 mm 2 wires for 

through circuits running from electrical 


51

connectors (ШР2, ШР3, ШР4) installed at 

the side connector board. These circuits 

(about 65 of them) can be used for 

spacecraft/ground equipment link. For this 

purpose, a new cable will be run from the 

specified connectors to the third umbilical 

connector of the spacecraft. 

If the characteristics (wire cross-section, 

shielding, etc.) of the above-mentioned 

standard LV electrical links are not 

sufficient to ensure the required electrical 

interface with the SC, an additional cable 

may be inserted into the electrical 

connectors (ШР2, ШР3, ШР4) of the side 

connector board. At the LV/SHM joint, the 

output connectors of this additional cable 

will be mated with the SHM cables that 

service the Spacecraft. 

It is desirable to install on the spacecraft 

several hermetically sealed subminiature 

plug connectors used on SS-18 (standard 

designation ОС РС 50ГБАТ ВЛО.364.046 

ТУ), the pins of which are soldered with 

0.12 – 0.35 mm 2 wires. The plug weight is 

38 grams (with the case) and 20 grams 

(without the case). Maximum amperage is 

75 A. This connector can withstand low 

pressure (down to 10 -6 mm of Mercury) 

and pressure difference of up to 1 

atmosphere. This connector has a 

standard SS-18 no-impact disconnection 

device. 

All electrical connectors used in through 

circuits are standard proven devices 

successfully tested by a great number of 

SS-18 launches. Other types of umbilical 

Figure 8.2-1 Schematics of SC/SC COE Transit Electrical Links 


52

connectors can be installed on the 

Spacecraft. Their counterparts are to be 

provided to ISC Kosmotras for fabrication 

of the appropriate SHM cables. 

When the SC is injected into the required 

orbit, the LV control system, if necessary, 

issues a reference command to the SC 

instrumentation. Upon 0.1 second 

following the issuance of the command, 

the electrical connectors with SC are 

disconnected and 0.5 second after the 

issuance of the command, the SC/LV 

separation occurs. 

If active electrical interface is used or it is 

necessary to monitor the SC status up to 

the moment of launch, the SC COE can be 

installed in underground Facility 25 located 

in close vicinity of the silo launcher. 

During electrical checks of the SC 

integrated with LV inside the silo, the 

following parameters of SC/LV interface 

are monitored: 

 

 

 

 

connection of cables running from LV 

to SC and ground equipment; 

issuance of signal “INITIAL STATUS 

OF LV CONTROL SYSTEM”; 

resistance of power lines supplying 

signals from the LV control system to 

the SC instrumentation; and 

issuance of reference signals from the 

LV control system to the SC 

instrumentation. 

The list of commands (signals) required for 

maintaining interface with the LV control 

system during LV preparation for launch, 

launch and flight, as well as sequence and 

parameters of such interface, types of 

connectors and number and 

characteristics of electrical links are 

agreed upon between ISC Kosmotras and 

the spacecraft authority. 


53

9. Spacecraft Environments 

 

1.3 during the LV flight. 

9.1 Stiffness Criteria (Frequency 

Requirements) 

The spacecraft should be designed with a 

structural stiffness, which ensures that the 

values of fundamental frequency of the 

spacecraft, hard mounted at the 

separation plane, are not less than: 

 

 

20 Hz in the longitudinal axis; and 

10 Hz in the lateral axis. 

If it is not possible to comply with the 

above requirements, SDB Yuzhnoye will 

carry out an additional analysis of the LV 

dynamic characteristics and loads that will 

take into account the spacecraft 

fundamental frequencies. 

9.2 Quasi-static and Dynamic Loads 

Tables 9.2-1 and 9.2-2 contain quasi-static 

and dynamic components of accelerations 

that act on the SC/LV interface during the 

ground handling, launch and in-flight. 

Spacecraft dimensioning and testing must 

take into account safety factors, which are 

defined by the spacecraft authority, but 

should be no less than the values given 

below: 

 

 

2.0 for ground handling; 

1.5 during launch while LV is moving 

inside the TLC; 

The spacecraft should remain operable 

after the effect of the above accelerations. 

9.3 Vibration Loads 

Described below are vibrations acting on 

the Spacecraft attachment points during 

the LV flight. Two types of vibrations are 

as follows: 

 

 

Harmonic oscillations; and 

Random vibrations. 

The harmonic oscillations are 

characterized by the amplitude of vibroaccelerations 

and frequency. The 

parameters of harmonic oscillations are 

given in Tables 9.3-1 and 9.3-2. 

The random vibrations are characterized 

by spectral density of vibro-accelerations 

and the duration of influence. The random 

vibration parameters are given in Table 

9.3-3. 

The random vibrations are spatial with 

approximately equal intensity of vibroaccelerations 

in each of the three 

randomly selected mutually perpendicular 

directions. 

The values of amplitude and spectral 

densities are given in the extreme octave 

points. The change of these values within 

the limits of each octave is linear in the 

logarithm frequency scale. 

 

1.3 during launch after the LV exits 

from the TLC; 


54

Table 9.2-1 Accelerations at SC/LV Interface during Transportation 

Load Source 

Acceleration 

Longitudinal (X) Lateral (y) Lateral (z) 

SHM Transportation ±0.4 -1.0±0.7 ±0.5 

Table 9.2-2 Maximum Quasi-static and Dynamic Accelerations at SC/LV Interface 

Load Source 

LV movement inside TLC 

After LV exit from TLC 

1 st stage burn: 

Maximum dynamic head 

Maximum longitudinal acceleration 

2 nd stage burn – maximum 

longitudinal acceleration 

Acceleration 

Longitudinal (X) Lateral (y, z) 

2.5±0.7 

±1.0 

3.0±0.5 

7.5±0.5 

±0.3 

±0.8 

0.5±0.5 

0.1±0.5 

7.8±0.5 0.2 

3 rd stage burn -0.3...-0.5 0.25 

Notes to Tables 9.2-1 and 9.2-2: 

 

 

 

 

Lateral accelerations may act in any direction, simultaneously with longitudinal 

ones; 

The above values are inclusive of gravity force component; 

Dynamic accelerations are preceded by "±" symbol; 

The above values are correct for the spacecraft complying with the fundamental 

frequency requirements contained in paragraph 9.1. 


55

Table 9.3-1 Amplitude of Harmonic Oscillations at SC/LV Interface. Longitudinal 

Axis (X) 

Frequency sub-band, Hz 5-10 10-15 15-20 

Amplitude, g 0.5 0.6 0.5 

Duration, sec. 10 30 60 

Table 9.3-2 Amplitude of Harmonic Oscillations at SC/LV Interface. Lateral Axes (Y, Z) 

Frequency sub-band, Hz 2-5 5-10 10-15 

Amplitude, g 0.2-0.5 0.5 0.5-1.0 

Duration, sec. 100 100 100 

Table 9.3-3 Spectral Density of Vibro-accelerations at SC/LV Interface 

Frequency sub-band, Hz 

Liftoff, LV flight 

segment where M=1, 

q max 

Load Source 

1 st stage burn (except for LV 

flight segment where М=1, 

q max ), 2 nd stage burn, 3 rd 

stage burn 

Spectral Density, g 2 /Hz 

20-40 0.007 0.007 

40-80 0.007 0.007 

80-160 0.007-0.022 0.007 

160-320 0.022-0.035 0.007-0.009 

320-640 0.035 0.009 

640-1280 0.035-0.017 0.009-0.0045 

1280-2000 0.017-0.005 0.0045 

Root Mean Square Value, σ, g 6.5 3.6 

Duration, sec. 35 831 


56

9.4 Shock Loads 

Shock loads are wide-band, fading 

processes and are characterized by the 

shock spectrum and the duration of action. 

The activation of the separation pyrodevices 

is a source of the vibro-pulse 

loads at the spacecraft attachment points 

(the duration of shock process is up to 

0.1 sec). The shock spectrum values are 

given in Table 9.4-1. They are accurate for 

the Q=10 and for each of the three 

randomly selected mutually perpendicular 

directions. The change of the shock 

spectrum values versus frequency within 

each sub-band is linear (in the logarithm 

frequency scale and shock spectrum 

values). 

9.5 Acoustic Loads 

The sources of acoustic loads are: 

 

 

1 st stage motor burn; 

frame surface pressure fluctuations in 

the turbulent boundary layer. 

The acoustic loads are characterized by 

the duration of action, integral level of the 

sound pressure within the frequency band 

of 20-8,000 Hz, and the levels of sound 

pressure within the octave frequency 

band with the mean geometric 

frequencies of 31.5; 63; 125;…; 2,000; 

4,000; 8,000 Hz. 

Table 9.4-1 Shock Spectrum at Spacecraft Attachment Points 

Load Source 

30- 

50 

50- 

100 

Frequency subband, Hz 

100- 

200 

200- 

500 

500- 

1000 

1000- 

2000 

2000- 

5000 

Number 

of shock 

impacts 

Separation of fairing, 3 rd 

stage and neighboring 

spacecraft 

5- 

10 

10- 

25 

Shock Spectrum Values, g 

25- 

100 

100- 

350 

350- 

1000 

1000 1000 * 

Separation of SC 

5- 

10 

10- 

25 

25- 

100 

100- 

350 

350- 

1000 

1000 

1000- 

3000 

1 

Note: * - number of shock impacts is contingent on a number of spacecraft installed in 

the SHM. 


57

Table 9.5-1 Acoustic Loads 

Mean Geometric Frequency of Octave Frequency 

band, Hz 

31.5 

63 

125 

250 

500 

1000 

2000 

4000 

8000 

Level of Sound Pressure, dB 

125 

132 

135 

134 

132 

129 

126 

121 

115 

Integral Level of Sound Pressure, dB 140 

Duration, sec. 35 

9.6 Temperature and Humidity 

Conditions and Thermal Effect on 

Spacecraft 

During operations with the spacecraft at 

SC processing facility, the air temperature 

around the spacecraft is maintained within 

21 - 27 0 C, with relative humidity of not 

more than 60%. 

During SC/SHM integration at AITB, the 

air temperature is maintained within 5 - 

35 0 C, with relative humidity of not more 

than 80%. 

When transporting the SHM to SHM 

processing facility and to the launch silo, 

the temperature inside the Transporter- 

Erector is within 10-25 0 C with relative 

humidity of no more than 80%. 

During operations with the SHM at SHM 

processing facility, the air temperature 

around the spacecraft is maintained within 

5 - 35 0 C, with relative humidity of not more 

than 80%. 

When loading the Space Head Module 

into the launch silo and mating it with the 

LV, the SHM is affected by the 

temperature within 0-45 0 C during the time 

period of no more than 30 minutes and 

with the temperature within 5 - 35 0 C during 

the time period of no more than 5.5 hours, 

with the relative humidity being no more 

than 80%. 

When the SHM is inside the silo, the 

temperature inside the silo is within the 

range of 5 - 25 0 C with the possible shortterm 

increase of up to 35 0 C and relative 

humidity is of no more than 80%, and the 

temperature around the spacecraft is 


58

within the range of 5 - 30 0 C with the 

relative humidity being no more than 70%. 

Spacecraft heat emission while on the LV 

inside the silo and in-flight were not taken 

into account. 

Thermal flux acting on the spacecraft from 

the inner surface of gas-dynamic shield 

will not exceed 1,000 Wt/m 2 . 

9.7 Pressure Underneath LV Fairing 

Pressure change inside the fairing 

envelope during the ascent phase is given 

in Figure 9.7-1. 

The maximum rate of in-flight pressure 

change inside the fairing envelope does 

not exceed 0.035 kgf/(cm 2 per sec.), 

except for transonic phase of flight where 

a short term (2-3 seconds) increase up to 

0.035 kgf/(cm 2 per sec.) is possible. 

Data contained in this section may be 

specified for each specific mission. 

9.8 Gas-dynamic Effect on Spacecraft 

Following separation from the Space Head 

Module the spacecraft encounters a short 

term impact (several seconds) of the 3 rd 

stage motor plume. 

All combustion products (composed of: N 2 

– 28%, H 2 – 27%, H 2 O – 21%, CO 2 – 18%, 

CO – 6%) are in gaseous state; solid or 

liquid phases are not present. 

Parameters of the 3 rd stage motor plume 

affecting the spacecraft are given in Table 

9.8-1. 

1.2 

1.0 

0.8 

0.6 

0.4 

0.2 

0.0 

-0.2 

P, kgf/cm 2 Figure 9.7-1 Pressure Change Rate inside Payload Envelope 

0 20 40 60 80 100 t,sec. 


59

Spacecraft surface contamination due to 

sedimentation of solid or liquid particles 

does not occur, since they are not present 

among the 3 rd stage motor combustion 

products. 

Spacecraft surface contamination due to 

H 2 O vapor condensation does not occur, 

since the maximum gas pressure on the 

spacecraft surface (stagnation pressure) is 

significantly (dozens of times) lower than 

H 2 O saturated vapor pressure at the 

spacecraft surface temperature, which, at 

this moment, normally exceeds 273 0 K. 

3 rd stage motor plume impact on the 

spacecraft is insignificant. Integral thermal 

flux on the spacecraft surface during 

separation will not exceed 5 Wt per 

hour/m 2 . Actual heating of 1 mm thickness 

spacecraft shell made of AMg-6 alloy is 2- 

4 0 . 

3 rd stage motor plume may cause light 

disturbances of the spacecraft motion. 

Torque may be induced that is dependant 

on the spacecraft position with respect to 

the X axis of the launch vehicle and 

spacecraft inertia characteristics. 

Data on the spacecraft launched by 

Dnepr-1 LV and history of their 

subsequent operation confirm the absence 

of 3 rd stage motor plume impact on 

spacecraft operability, including the 

spacecraft equipped with high resolution 

optics. 

9.9 SC/LV Electromagnetic Compatibility 

The engineering solution to use EPM or 

gas-dynamic shield provides for about 30 

dB noise shielding within the frequency 

range of 10 kHz – 30 GHz. Apart form 

that, up to about 275 th second of flight, the 

LV fairing provides for additional 10 – 20 

dB noise shielding within the frequency 

range of 10 kHz – 1000 MHz. 

Maximum residual levels of LV on-board 

system (radioelectronics) RF emissions 

that penetrate the EMP (GDS) and have 

effect on the spacecraft during the active 

phase of LV flight prior to and after the 

EMP cover (GDS) separation, are given in 

Table 9.9-1. 

Table 9.8-1 Maximum Values of 3 rd Stage Motor Plume Parameters Affecting Spacecraft 

х, m 0 1.8 3.3 4.5 6 7.5 9 15 20 30 50 

t,sec. 0 1.2 1.6 1.9 2.2 2.5 2.7 3.5 4.0 4.9 6.3 

Р 0 , kgf/m 2 0.2 6.27 4.7 7.3 11 8.2 6.6 3.2 1.84 0.84 0.3 

Р, kgf/m 2 0.05 0.048 0.034 0.025 0.017 0.012 0.0095 0.0038 0.002 0.00077 0.00023 

Т, К 254 252 238 226 213 200 193 165 149 127 104 

V, m/sec. 3670 3671 3680 3687 3696 3704 3709 3726 3736 3750 3764 

Where: x - distance (along X axis) from EPM attachment plane; 

P 0 - gas stagnation pressure; 

P, T, V - pressure, temperature and velocity respectively of undisturbed gas flow. 


60

Maximum levels of industrial noise 

induced by LV control system equipment 

that affect the spacecraft during active 

phase of LV flight within the frequency 

range of 10 kHz – 1 GHz, will not exceed 

5.6 mV prior to and 35 mV after the EMP 

cover (GDS) separation. 

Maximum levels of electromagnetic 

emissions (radio and industrial noise) for 

LV on-board systems are given in Table 

9.9-2. 

Maximum levels of electromagnetic 

emissions generated by spacecraft and 

Table 9.9-1 Maximum Residual Levels of RF Emissions 

Frequency Band 

10 kHz – 140.4 MHz 

140.4 – 144.4 MHz 

144.4 – 1000.5 MHz 

1000.5 – 1004.5 MHz 

1004.5 – 2500 MHz 

2500 – 2855 MHz 

2855 – 2865 MHz 

2865 – 30000 MHz 

Electormagnetic Field Strength, V/m 

Prior to EPM (GDS) Cover 

Drop 

1.4·10 -2 

2.2·10 -2 

0.43 

6.9 

1.4·10 -2 

2.2·10 -2 

0.38 

6 

1.4·10 -2 

2.2·10 -2 

5.4·10 -2 

0.12 

1.2 

19 

5.4·10 -2 0.12 

After EPM (GDS) Cover 

Drop 

Table 9.9-2 Maximum Levels of Electromagnetic Emissions 

Frequency Band 

10 kHz –125 MHz 

125 – 250 MHz 

250 – 1000 MHz 

1000 – 1050 MHz 

1050 – 1570 MHz 

1570 – 1620 MHz 

1620 – 2750 MHz 

2750 – 2900 MHz 

2900 – 7500 MHz 

7500 – 7600 MHz 

7600 – 30000 MHz 

Electormagnetic Field Strength 

70 mV/m 

10 V/m 

70 mV/m 

10 V/m 

70 mV/m 

10 µV /m 

70 mV/m 

50 V/m 

70 mV/m 

10 V/m 

70 mV/m 


61

spacecraft electric checkout equipment 

(ECOE) at the SC/LV and LV/spacecraft 

ECOE interfaces (at a distance of 1 m 

from spacecraft and its ECOE), from the 

beginning of SC/LV integration and until 

spacecraft separation from LV plus 1 

minute, must be 10 dB less than the 

values given in Table 9.9-2. 

During preparation and launch of 

spacecraft that has active electrical 

interface with LV, the spacecraft 

instrumentation is also affected, along the 

electrical circuits of SC/LV interface, by 

electromagnetic interference with the 

levels of up to 100 dB/µV within the 

frequency range of 30 Hz – 100 MHz for 

feed and control circuits, and up to 60 

dB/µV within the frequency range of 30 Hz 

– 30 GHz – for data circuits. 

Instrumentation of SC, in case of active 

electrical interface, should not generate 

electromagnetic emissions in the electric 

circuits of the SC/LV interface over 60 

dB/µV within the frequency range of 30 Hz 

– 100 MHz – for feed and control circuits, 

and over 40 dB/µV within the frequency 

range of 30 Hz – 30 GHz – for data 

circuits. 

Density of electromagnetic fields 

generated by Cosmodrome electronic 

equipment at processing facilities and 

routes of transportation of SC, SHM and 

LV do not exceed 10 V/m for frequency 

range 10 kHz –30 GHz, except for the 

frequency range of 1570 - 1620 MHz, 

where maximum allowable level of 

external electromagnetic interference is 30 

mV/m (for active or shut radio receivers of 

spacecraft). This is similar to both 

European and US standards. 

Coordination of efforts to ensure 

electromagnetic compatibility of SC, LV 

and range systems at all stages of SC/LV 

integration is done under a special EMC 

procedure for a specific mission that is 

developed at the preliminary integration 

phase. 

9.10 Spacecraft Tests Required to Meet 

Dnepr LV Launch Services Requirements 

The customer shall demonstrate that the 

spacecraft meets the requirements 

detailed in the entire section 9 of this 

User’s Guide, by means of analyses and 

ground tests. 

For spacecraft qualification and 

acceptance, sinusoidal, shock and random 

tests are mandatory. 

A test plan established by the spacecraft 

authority describing the tests, which are 

executed on the spacecraft, shall be 

provided to SDB Yuzhnoye. 

After completion of the tests, the test 

results report shall be submitted to SDB 

Yuzhnoye. 

10. Ground Qualification Tests 

To verify the ability to integrate the 

spacecraft with the Dnepr launch vehicle 

and to confirm the operability of 

spacecraft/launch vehicle mechanical and 

electrical interfaces, a ground qualification 

test program may be provided that 

includes spacecraft fit-check testing, 

vibration testing of spacecraft and SHM 

structural elements and spacecraft 

separation system tests. 

The objective of qualification testing is to 

confirm the following: 


62

operations with spacecraft provided for 

by spacecraft authority are easy to 

handle by personnel and ensure nocollision 

integration with launch vehicle; 

engineering solutions to protect the 

spacecraft from damage are sufficient; 

vibration strength at spacecraft 

attachment points is sufficient to 

withstand loads on the spacecraft 

during LV launch and flight; 

 

separation system remains operable 

following the impact of vibration loads; 

complete separation along 

spacecraft/adapter joints is achieved; 

 

motion parameters of the spacecraft 

being separated meet the 

requirements and shock loads that act 

upon activation of spacecraft 

separation pyros are within the 

allowable limits. 

Qualification tests are performed at SDB 

Yuzhnoye facilities with participation of 

spacecraft authority specialists. The scope 

of qualification testing is to be agreed 

upon with the spacecraft authority. 

Necessary process related equipment and 

spacecraft dummies with actual 

mechanical and electrical interfaces to be 

furnished by spacecraft authority are used 

for qualification testing. Based on the 

results of qualification testing, an 

acceptance certificate authorizing the 

launch of a specific spacecraft on Dnepr 

LV is released. 


63

11. Telemetry and Tracking 

Telemetry system is based on LV onboard 

radio telemetric system with data 

capacity of 512 Kbit/sec., which allows to 

register both slowly and rapidly changing 

parameters. 

During spacecraft injection, a number of 

physical process parameters are 

registered, which determine the proper 

functioning of the LV systems and units, 

characterize the LV/SC separation 

process as well as measure dynamic and 

heat impact on the spacecraft during its 

orbital injection. For this purpose, the LV 

on-board telemetry system is equipped 

with initial signal transformers and 

normalizers. 

Data capacity of the LV telemetry system 

allows, if necessary, for recording 

information supplied by the spacecraft onboard 

instrumentation within the broad 

sampling frequency range of 25 Hz – 8 

kHz and output voltage of 0-6 V. 

Motion parameters of the LV center of 

gravity are determined by GPS system, 

which transmits obtained data through 

telemetry channels. 

When launching spacecraft on Dnepr LV 

from Baikonur eastward (orbit inclinations 

50.5 and 64.5 degrees), the telemetry data 

transmitted from LV through two channels 

(to ensure appropriate transmission 

reliability) is registered by Baikonur ground 

tracking stations and ground tracking 

stations of Russian Federation located 

along the LV flight trajectory. 

When launching spacecraft from Baikonur 

southward (orbit inclinations 87.3 and 98.0 

degrees), the telemetry data is 

subsequently registered by a mobile 

ground tracking station located in Oman 

near the town of Salalah. This ground 

station relays to Baikonur data collection 

and processing center a certain amount of 

telemetry data in real-time scale, including 

the parameters of spacecraft initial motion. 

Specific requirements for telemetry 

reading and interface will be defined 

during SC/LV integration concept work. 

Pre-launch checks of the LV on-board 

telemetry system are carried out by means 

of available standard ground checkout 

equipment. 

Figure 11-1 Tracking Station 


64

12. Analysis of Flight Results 

Analyses of all LV system functioning, of 

the loads the spacecraft encountered 

during injection are carried out based on 

the results of the telemetry measurements 

taken during the LV pre-launch 

preparation and flight. 

The scope of analyses required is 

determined jointly with the customer in the 

process of approval of the statement of 

work for the spacecraft launch. 

The analysis of the flight results is carried 

out in the following way: 

 

 

 

preliminary data on spacecraft initial 

motion parameters (2 hours after 

spacecraft separation); 

full set of data on spacecraft orbital 

parameters (1 day after spacecraft 

separation); 

comprehensive analysis of all the data 

received in order to evaluate the LV 

system and sub-system functioning, to 

get their quantitative values, to assess 

spacecraft injection environments and 

to determine the probable reasons of 

malfunctions if there were any (1 

month after launch). 

 

 

real-time telemetry data on a limited 

number of LV system status 

parameters (command issuance and 

execution, propulsion unit operation 

and LV flight stabilization); 

determination of spacecraft separation 

event (10 minutes after the separation); 

determination of spacecraft actual 

separation time (1.5 hours after the 

spacecraft separation event); 

Figure 12-1 Flight Data Processing 

Center 


65

13. Range Safety 

 

effect of chemicals and pollutants; 

ISC Kosmotras possesses certain 

expertise and proven safety procedures, 

both organizational and technical, for 

launch campaigns. All works and 

operations are conducted in strict 

accordance with approved plans and 

documents. 

Our company guarantees that the 

Customer’s representatives have full 

control over the access to the spacecraft, 

as far as it is technically feasible. Joint 

operations of Customer’s and Provider’s 

specialists are conducted in compliance 

with the approved procedures. Following 

the completion of operations in clean room 

facility (SC/adapter integration, assembly 

of EPM), the EPM cover attachment points 

are sealed by Customer’s representatives, 

this fact is registered in a special 

certificate and recorded by video and 

photography. Following this procedure, the 

direct access to the spacecraft becomes 

impossible. If necessary, visual and 

remote monitoring of the EPM (gas 

dynamic shield, in case of a large 

spacecraft) and the entire space head 

module seal points can be arranged. 

For all operations with the spacecraft, 

safety procedures are worked out, planned 

and followed with regard to the following 

hazards listed below: 

 

 

electrical hazards; 

explosion hazards; 

 

 

 

mechanical impact; 

weather, thermal and light effects; 

space debris; 

personnel mishaps and equipment 

failure; 

 

acts of God (earthquakes, hurricanes, 

showers). 

ISC Kosmotras is responsible for ensuring 

safety during work with spacecraft, ground 

support equipment at all stages of 

spacecraft preparation for launch, in full 

compliance with the standards and 

requirements of the Russian Federation 

and spacecraft authority. 

Spacecraft authority is responsible for 

compliance with the specific safety 

requirements during development, 

fabrication and operation of the spacecraft 

and ground support equipment, as well as 

during its personnel stay at the 

Cosmodrome. 

Spacecraft authority submits to ISC 

Kosmotras the documents that describe: 

 

 

hazardous systems and operations of 

the spacecraft and its support 

equipment; 

results of hazardous systems and 

elements testing; and 

 

 

 

fire hazards; 

UHF emission hazards; 

biological hazards; 

results of spacecraft tests for 

vibrations and flight loads. 


66

ISC Kosmotras reviews these documents 

and has the right to issue its comments for 

each of such safety documents. 

If necessary, ISC Kosmotras is prepared 

to develop, together with the spacecraft 

authority, and sign a comprehensive 

document that governs the safety 

procedures and requirements for a specific 

spacecraft. 

The following safety precautions are taken 

during operations with spacecraft 

(tentative list): 

General: 

 

all equipment used for operations with 

the spacecraft, space head module 

and launch vehicle is certified as per 

the effective standards for the period of 

time sufficient to successfully complete 

all operations; 

required temperature and humidity 

conditions are maintained throughout 

al stages of operation; 

 

only authorized personnel have access 

to the spacecraft processing area; 

uninterrupted power supply is 

guaranteed during spacecraft 

integration and checkout; 

 

 

required speed limits are complied with 

during spacecraft and ground support 

equipment transportation; 

three levels of supervision are ensured 

for the control of critical and most 

hazardous operations with the space 

head module and launch vehicle. 

During handling: 

 

 

 

 

Container off-loading is done by lifting 

gear of the airplane, by road cranes or 

forklifts using pull ropes, cross-arms, 

pallets or other gear of the spacecraft 

authority; 

Spacecraft offloading is conducted in 

strict accordance with operational 

documents; 

Only authorized personnel are allowed 

on operations site. Operations site 

must be sufficiently lit up and, if 

possible, fenced off; 

During offloading, spacecraft must be 

tied by means of halyards to prevent 

rocking. 

During transportation: 

 

 

Arrived containers with the spacecraft 

and ground support equipment will be 

delivered to the processing facility for 

subsequent operations by road or rail, 

on common carrier rail flatcars in 

shrouded condition to protect from 

direct impact of precipitation; 

If necessary, the containers can be 

transported from Yubileiniy airport on 

special environmentally controlled 

railcars; 

At the request of the spacecraft 

authority, the ground transport loads 

(g-loads, accelerations, temperature) 

can be controlled during transportation 

and the results made available to the 

spacecraft authority. 

During operations at spacecraft and space 

head module processing facility: 

 

spark-proof tools must be used. The 

personnel working with the spacecraft 


67

must have special gear for static 

electricity elimination; 

Checkout equipment must be 

grounded and resistance of bonding 

connection must be checked; 

 

 

Personnel must be warned prior to 

supplying power or pressure to the 

spacecraft systems; 

Personnel must be warned prior to 

activating on-board transmitting units; 

When integrating the spacecraft 

elements covered with thermal 

insulation, the personnel must wear 

respirators and cotton gloves. Upon 

completion of work, the clothes must 

be de-dusted and open skin washed 

with water and soap; 

 

Personnel must wear special clothes 

and soft shoes, and comply with 

operational 

documentation 

requirements. 

Safety during the operations with the 

space head module, space head 

module/launch vehicle integration and 

checkout, is ensured through exercising 

the technical, organizational and design 

measures listed below: 

 

 

launch vehicle checkout sequence is 

fully automated that precludes 

emergencies due to operator’s errors; 

prior to operations with the launch 

vehicle, the following checks are 

performed to avoid improper electrical 

connection with actuators and damage 

to the launch vehicle on-board 

equipment and cable network: 

availability of electrical connection in 

circuits, absence of power line bridging 

and connection with the launch vehicle 

frame, insulation resistance, initial 

condition of equipment, proper mating 

of connectors, integrity of pyro-device 

fuse heads, in order to prevent 

improper supply of power to actuators 

and avoid damage to the launch 

vehicle on-board equipment and 

harness; 

use of connectors with different 

configuration and cables of different 

length to avoid improper connection; 

 

use of 30 V power line from the short 

circuit protected ground power sources 

to the launch vehicle on-board 

equipment; 

use of such on-board equipment 

design that provides for proper 

grounding to avoid accumulation of 

static electricity on launch vehicle and 

space head module frame; 

 

monitoring that the power supplied to 

the pyro-devices and launch vehicle 

frame during the checks of launch 

vehicle and space head module, has 

the proper parameters. 

Safety during operations with the space 

head module, its integration with the 

launch vehicle and checks of the launch 

vehicle integrated with the space head 

module, is ensured through the application 

of the procedures provided for in the 

operational documentation and 

compliance with the Russian Federation 

requirements “OTT 1.1.10 –90, OTT 

11.1.4 –88, part 6, GOST B 20.39.308 – 

76. Effectiveness of these procedures was 

confirmed by the results of the fit-check 

and bench testing conducted at SDB 

Yuzhnoye facilities and by the successful 

launches in 1999 and 2000. 


68

14. Transportation of Spacecraft and 

Associated Equipment. Customs 

Clearance 

Spacecraft authority ensures the delivery 

of dummy spacecraft and associated 

equipment to Ukraine for ground testing 

and of spacecraft and associated 

equipment to Baikonur Cosmodrome, and 

obtains their customs clearance. 

All operations with the spacecraft, 

equipment and dummy spacecraft are 

divided into 4 stages: 

1. Transportation of dummy spacecraft 

and associated equipment to Ukraine 

for ground testing (Customer’s country 

– SDB Yuzhnoye, Dnepropetrovsk, 

Ukraine) and their subsequent return, 

upon completion of ground testing, to 

customer’s country; 

2. Customs clearance of dummy 

spacecraft and equipment upon entry 

into Ukraine (SDB Yuzhnoye, 

Dnepropetrovsk). Customs clearance 

of dummy spacecraft and equipment 

upon exit from Ukraine; 

3. Transportation of spacecraft and 

associated equipment to the launch 

site (Customer’s country – Russia, 

Moscow – Baikonur Cosmodrome, 

Kazakhstan). Return of associated 

equipment to customer’s country after 

launch; 

4. Customs clearance of spacecraft and 

associated equipment upon entry into 

Russia (Moscow) and Kazakhstan 

(Baikonur Cosmodrome). Customs 

clearance of associated equipment 

upon exit from Kazakhstan (Baikonur 

Cosmodrome) and Russia (Moscow) 

after launch en-route to customer’s 

country. 

To accomplish stages 1, 3 and 4, ISC 

Kosmotras can recommend to conclude a 

contract with a Russian customs clearance 

and transportation broker company 

specializing in space related area to 

ensure import of space related equipment 

into Ukraine, Russia and Kazakhstan 

(Baikonur Cosmodrome). 

Accomplishment of Stage 2 can be 

organized by SDB Yuzhnoye. 

14.1 Transportation of Dummy Spacecraft 

to Ukraine for Ground Testing 

To ensure the entry of the dummy 

spacecraft and associated equipment into 

Ukraine for ground testing, it will be 

necessary for customer to conclude a 

formal contract with SDB Yuzhnoye. 

Based on this contract, dummy spacecraft 

will enter Ukraine. Transportation of the 

dummy spacecraft to Ukraine and their 

customs clearance may be entrusted to a 

Russian transportation and customs 

clearance company on a separate 

contract, or done by the dummy spacecraft 

owner independently. Upon completion of 

ground testing, the dummy spacecraft and 

associated equipment will be returned to 

the customer. Since this is a temporary 

import regime, the customs duties will be 

minimum. 

14.2 Transportation of Spacecraft and 

Associated Equipment to Baikonur 

Cosmodrome 

The most appropriate and proven scenario 

of transportation of spacecraft and 

equipment is the following: 


69

Customer concludes a direct contract 

with a Russian transportation and 

customs clearance broker company 

that delivers the spacecraft and 

equipment by regular or charter flight to 

Moscow, Russia (Sheremet’yevo-2 

International Airport); 

At Sheremet’yevo-2 the broker 

company obtains customs clearance 

for the spacecraft and equipment and 

then ships them by air to Baikonur; 

 

 

Upon arrival at a Baikonur airfield, the 

broker company obtains Kazakhstan 

customs clearance for the spacecraft 

and equipment; 

Then the spacecraft and equipment are 

 

transported to the spacecraft 

processing facility. Reloading and 

transportation of the spacecraft and 

equipment at Baikonur is supported by 

ISC Kosmotras; 

Return of the associated equipment 

after launch is done in reverse 

sequence. 

Duration of transportation from Europe is 

no more than 2 days. 

During the performance of all the abovementioned 

stages, ISC Kosmotras verifies 

the proper conduct of transportation and 

customs clearance in Russia, Kazakhstan 

(Baikonur Cosmodrome) and Ukraine. 


70

15. Principles of Pricing Policy and 

Contractual Relations 

ISC Kosmotras is a primary contractor 

responsible for all technical and 

organizational issues associated with the 

customer’s payload launch. 

Dnepr-1 LV is created on the basis of the 

converted SS-18 ICBM by minor 

modifications. At present, a sufficient 

number of SS-18s on storage is available. 

This fact is reflected in the price structure, 

which does not have the ”cost of LV 

fabrication” component, and therefore 

makes it possible to offer customers lower 

level of launch services price than that of 

other launch services providers. The lower 

price of launch services, reliability and 

proofness of all Dnepr SLS systems make 

a very attractive combination for potential 

customers. 

Dnepr-1 launch services price is 

contingent on the following factors: 

orbit parameters (type, altitude, 

inclination); 

 

 

weight and dimensions of spacecraft, 

spacecraft layout diagram; 

type of interface; 

spacecraft processing requirements 

(filling with propellant, high cleanliness 

requirements, etc.); 

 

type of launch: dedicated or cluster 

(consolidated), single or multiple 

(under a program of launches). 

To launch a spacecraft on Dnepr-1 LV, 

customer and ISC Kosmotras conclude a 

Launch Services Agreement (LSA). LSA 

for launch of a large spacecraft is 

concluded no less than 18 months prior to 

the planned launch date. LSA for launch of 

a small spacecraft may be concluded 10 

months prior to the planned launch date. 

The LSA consists of the following: 

 

Main part; 

Annex 1 – Launch Services 

Specifications; 

 

 

Annex 2 – Logistical Support of Work; 

Annex 3 – Organizational Issues. 

All annexes should be agreed upon by the 

parties before the contract signature. 

The following services to be rendered by 

ISC Kosmotras are normally included into 

the contract price: 

spacecraft adaptation to launch 

vehicle, including necessary interfaces, 

development, manufacturing and 

testing of LV flight adapter and other 

necessary equipment; 

 

 

 

 

provision of facilities and required 

conditions for spacecraft processing 

and offices for customer’s personnel at 

the Cosmodrome; 

launch of spacecraft into the required 

orbit and supply of information on 

spacecraft separation from LV, on 

orbital parameters at the moment of 

separation; 

third party liability insurance; 

performance of and payment for the 

transportation of the spacecraft and 


71

equipment from a Baikonur 

Cosmodrome airport to the spacecraft 

processing facility, and for the 

transportation of the equipment back to 

that airport. 

Under the LSA, project managers are 

appointed by the customer and ISC 

Kosmotras. They are responsible for 

organizational and coordination efforts 

throughout the effective period of the LSA. 


72

16. Design and Technical Documents to 

Be Submitted by Spacecraft Authority 

To evaluate the possibility of using Dnepr 

Space Launch System for launch of a 

specific spacecraft as well as to prepare 

the materials for spacecraft/Dnepr SLS 

interface, the spacecraft authority shall 

submit the following materials: 

1) Line and dimensional drawings of 

spacecraft (including its adapter); 

2) Spacecraft envelope drawing (The 

envelope where the spacecraft and 

adapter are placed taking into 

account all protruding elements and 

possible deviations from the 

nominal position due to errors 

during its manufacturing and 

balancing as well as deformations 

during its joint operation with the LV 

with respect to the plane of the 

adapter interface with the LV. The 

axis of the spacecraft envelope will 

be a perpendicular to the plane of 

the adapter interface with the LV 

drawn from the center of the 

circumference on which the 

attachment holes are located); 

3) Detailed drawing of the spacecraft 

(adapter) interface with the LV; 

4) Detailed drawing of the spacecraft 

interface with adapter denoting the 

elements of the separation system 

(separation mechanisms, guiding 

pins, attachment elements, etc.) 

and their characteristics with 

tolerances and deviations). The 

drawing that shows the area of 

spacecraft structural elements 

penetration into the adapter 

denoting structural gaps and 

“dangerous” points; 

5) The list of connection lines to the 

adapter describing their 

characteristics (the number and 

type of connection lines, their 

location, disconnection force and 

travel with tolerances and 

deviations); 

6) The diagram showing the location 

of the separation telemetry sensors 

at the spacecraft/adapter interface 

and their characteristics 

(disconnection force and travel with 

tolerances and deviations); 

7) Mass, Center of Gravity and inertia 

characteristics of the spacecraft, 

adapter as well as the 

spacecraft/adapter assembly with 

tolerances and deviations for all 

variants and modifications of the 

spacecraft, and also the distribution 

of weight and linear moment of 

inertia along the length of the 

spacecraft/adapter assembly; 

8) Stiffness characteristics of the 

spacecraft (the change of bending 

rigidity EJ, longitudinal rigidity EF, 

and torsion rigidity GJp along the 

length of the spacecraft/adapter 

assembly); 

9) Data on spacecraft structural 

elements separated in flight, 

requirements and constraints 

regarding the separation time; 

10) Parameters of the initial orbit in the 

injection point: inclination, apogee 

and perigee altitude with respect to 

earth mean radius (R m = 6,371 km), 

argument of perigee latitude, 

ascending node longitude; 

11) Injection accuracy requirements; 


73

12) Spacecraft longitudinal load 

constraints; 

13) Constraints for dynamic pressure at 

the moment of fairing separation; 

14) Spacecraft injection time 

constraints; 

15) The number, type and electric 

characteristics of the spacecraft 

separation system pyros; 

16) The list of the LV control system 

commands used by spacecraft, type 

and electric characteristics of the 

load on the LV control system 

circuits from the spacecraft; 

17) The diagram of interface of the 

spacecraft on-board equipment and 

the LV control system: 

diagram of the spacecraft 

separation system pyro connection; 

 

 

diagram of receiving commands 

from the LV control system; 

diagram of connectors along the 

LV-spacecraft (adapter) joint; 

18) The list and characteristics of the 

spacecraft links with ground 

equipment; 

19) Schedule of processing the 

spacecraft at the launch complex 

specifying the exchange signals 

with the LV control systems 

proposed by the spacecraft 

authority; 

20) Spacecraft processing flow chart; 

21) Safety requirements for spacecraft 

operation with the LV; 

22) Materials describing the volume of 

work with the spacecraft at the 

processing facility and the launch 

complex in case of launch 

cancellation; 

23) Initial data for accommodation of 

the spacecraft associated 

equipment at processing facility and 

launch complex premises; 

24) Spacecraft temperature and 

humidity control requirements at 

processing facility and launch 

complex; 

25) Initial data for spacecraft injection 

timeline; 

26) The list operations conducted on 

spacecraft during the phase of its 

joint flight with the LV; 

27) The list of telemetry and trajectory 

measurement devices and radio 

frequencies used on-board the 

spacecraft as well as sequence of 

their operation; 

28) Requirements to telemetry readings 

to be taken at the spacecraft (type 

and number of sensors, signal 

characteristics, registration time 

etc.); 

29) Documents confirming the 

completion of the spacecraft tests 

for static and dynamic loads with 

the level of loading exceeding the 

actual loads acting on the 

spacecraft during its flight with LV. 

The above list of documents may be 

amended and specified during joint work 

for detailed integration of the spacecraft 

and the Dnepr LV. 


74

17. Documents to Be Submitted to ISC 

Kosmotras by Spacecraft Authority 

Spacecraft mission nature and 

principal specifications; 

reside after it is placed into orbit stating 

that the spacecraft to be launched by 

the Russian Launch Vehicle will be 

registered in the national register of 

space objects; 

 

Spacecraft safety document to include 

information on fire-safety and explosion 

proofness; 

 

List of equipment temporarily imported 

by Customer for the launch, to be 

supplied to launch services Provider; 

 

 

 

Spacecraft nuclear safety document (in 

case of presence on-board spacecraft 

of nuclear power unit or radiation 

sources); 

Statement that the spacecraft will not 

be for military purposes from the 

appropriate government organization 

(e.g., the national space agency) in the 

country where ownership of the 

spacecraft will reside after it is placed 

into orbit; 

Written statement from the appropriate 

government organization (e.g., the 

national space agency) in the country 

where ownership of the spacecraft will 

Document on spacecraft radio 

frequency bands and maximum levels 

of UHF emission; 

 

 

Document from the owner of the 

spacecraft on its insurance aspects; 

and 

Spacecraft launch clearance document 

(after spacecraft integration to the 

launch vehicle payload module). This 

document shall be signed at Baikonur 

by Customer’s authorized 

representatives and presented to 

launch services Provider at the point of 

Change of Custody to clear the 

spacecraft for launch. 


75

18. Launch Campaign Schedule 

Figure 18-1 shows a launch campaign 

schedule for a large spacecraft. 

In case of a cluster launch, the time period 

between the contract signature and launch 

date may be reduced. 

# Milestone 

1 Signature of MoU between ISC 

Kosmotras and launch customer 

2 Project feasibility study 

3 Contract signature 

4 ICD release 

5 Release of documentation for additional 

hardware 

6 Fabrication of necessary hardware 

7 Fit-check 

8 Ground tests of payload module 

containing spacecraft dummies 

9 Delivery of launch vehicle, hardware and 

equipment to Baikonur 

10 Preparation and processing at Baikonur 

11 LV launch 

12 Preparation and release of launch result 

analysis data 

Quarter 

Q1 Q2 Q3 Q4 Q1 Q2 Q3 Q4 

Figure 18-1 Launch Campaign Schedule 


76

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